Curved cooling channels for a turbine component

Abstract

Turbine component, which includes: a blade (24) with a base (34) and a tip (36) arranged in the opposite direction to the base (34), wherein the blade (24) further comprises a pressure side (38) and a suction side (40) extending between a leading edge (42) and a trailing edge (44); a cooling channel (46) for the blade (24), which is at least partially arranged in the blade (24) and is designed to guide a coolant through the blade (24), and a curved cooling passage (60, 62) formed in the airfoil (24), wherein the curved cooling passage (60, 62) is flow-connected to the cooling circuit channel (46) of the airfoil (24) in such a way that the coolant flowing through the cooling channel (46) of the airfoil (24) is directed into the curved cooling passage (60, 62), wherein the curved cooling passage (60, 62) extends longitudinally in the airfoil (24) between the leading and trailing edges (42, 44) and at least partially along the pressure side (38) or the suction side (40) of the airfoil (24) between an upstream end (68) and a downstream end (70) of the curved cooling passage (60, 62), which terminate at an outer surface of the airfoil (24).

Description

FIELD OF THE INVENTION

[0001] The present subject matter relates generally to the cooling of turbine components, and more particularly to a turbine component having curved cooling passages for directing a coolant through the airfoil of the component. BACKGROUND OF THE INVENTION

[0002] In a gas turbine, hot combustion gases flow from an annular array of combustion chambers through a transition piece and through an annular hot-gas path. The turbine stages are typically distributed along the hot-gas path so that the hot combustion gases from the transition piece flow through the vanes and blades of the first stage and then through the vanes and blades of subsequent stages. The turbine blades may be attached to multiple impellers that form the turbine rotor, with each impeller attached to the rotor shaft so that the impeller rotates with the shaft.

[0003] A turbine blade generally comprises an airfoil extending radially upward from a substantially planar platform and a hollow shank portion extending radially downward from the platform. The shank portion may include a dovetail or other means to attach the blade to a wheel of the turbine rotor. Generally, during operation of a gas turbine, the hot combustion gases exiting the combustion chambers flow substantially over and around the airfoil of the blade. To protect the component from high temperatures, an airfoil of a blade typically includes a cooling passage used to supply the entire airfoil with a coolant, such as air, to reduce the temperature differential between the pressure and suction sides of the airfoil. The airfoil may also have a cooling scheme ora cooling arrangement for the air supply to the tip of the blade must be present.

[0004] Currently, rotor blade tips are cooled using a series of straight film cooling holes formed around the periphery of the airfoil and near the airfoil tip. These film cooling holes are drilled radially downward and inward from the airfoil surface into the airfoil cooling channel near the tip, allowing air flowing through the cooling channel to be directed upward into the tip. However, it has been found that this cooling arrangement does not provide optimal cooling for the blade tip. Furthermore, such a cooling arrangement requires drilling too many film cooling holes and consumes too much air. US Pat. No. 6,916,150 B2 describes a turbine blade for a turbine engine with a cooling system. The cooling system comprises one or more swirl chambers in the tip section of the turbine blade.The swirl chambers receive cooling fluid from metering slots that form a path between the blade's internal cooling cavities and the swirl chambers. The cooling fluids can be discharged through one or more film cooling openings. Furthermore, US Pat. No. 6,164,912 A provides a hollow airfoil having an outer wall and an internal cavity, wherein a plurality of cooling openings are arranged in the outer wall and are arranged spanwise along the leading edge of the airfoil. Furthermore, US Pat. No. 7,717,675 B1 describes a guide vane for use in a gas turbine engine, in which the guide vane has a plurality of near-wall mini-serpentine cooling modules arranged on the blade walls.

[0005] Accordingly, it would be desirable to provide a cooling arrangement for a blade of a turbine component that effectively cools sections of the blade, such as the tip, with fewer cooling openings and lower air consumption. BRIEF DESCRIPTION OF THE INVENTION

[0006] The object of the invention is achieved by the turbine component according to claim 1 and the turbine component according to claim 9. Embodiments and advantages of the invention will be set forth in part in the description which follows; they may also be obvious from the description or may be learned by practice of the invention.

[0007] According to a first embodiment, the turbine component comprises an airfoil, a cooling channel, and a curved cooling passage. The airfoil has a base and an oppositely arranged tip. The airfoil further comprises a pressure side and a suction side extending between a leading edge and a trailing edge. A cooling channel for the airfoil is at least partially arranged in the airfoil and configured to conduct a coolant through the airfoil. The curved cooling passage is formed in the airfoil and is substantially fluidly connected to the cooling channel of the airfoil such that the coolant flowing through the cooling channel of the airfoil is conducted into the cooling passage.The curved cooling passage also extends in the airfoil substantially longitudinally between the leading and trailing edges and at least partially along the pressure or suction side of the airfoil between an upstream end and a downstream end of the curved cooling passage terminating at an outer surface of the airfoil.

[0008] According to another embodiment, the turbine component comprises an airfoil, a cooling channel, and a plurality of curved cooling passages. The airfoil has a base and an oppositely disposed tip. The airfoil further comprises a pressure side and a suction side extending between a leading edge and a trailing edge. The airfoil cooling channel is at least partially disposed within the airfoil and configured to conduct a coolant through the airfoil. The plurality of curved cooling passages are formed in the airfoil, substantially proximate the airfoil tip. These cooling passages are fluidly connected to the airfoil cooling channel such that coolant flowing through the airfoil cooling channel is directed into the cooling passages.Each curved cooling passage extends substantially longitudinally in the airfoil between the leading and trailing edges and at least partially along the pressure side or the suction side of the airfoil between an upstream end and a downstream end of the curved cooling passage terminating at an outer surface of the airfoil.

[0009] These and other features, embodiments, and advantages of the present invention will become more fully understood when considered in conjunction with the following description and the appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0010] The following description presents a complete and practical disclosure of the present invention, including the preferred (best) embodiment. Reference is made to the following attached drawings: Fig. 1 shows a schematic representation of a gas turbine; Fig. 2 shows a perspective view of an embodiment of a turbine blade; Fig. 3 shows a partial perspective view of an embodiment of a turbine blade airfoil having a plurality of curved cooling passages formed therein in accordance with aspects of the present subject matter; Fig. 4 is a supervision of the Fig. 3 shown blade; Fig. 5 shows a partial perspective view of an example of a turbine blade airfoil having a plurality of curved cooling passages formed therein; Fig. 6 shows a top view of the Fig. 5 shown blade; Fig. 7 shows a partial side view of an example of a turbine blade airfoil having a plurality of curved cooling passages formed therein; Fig. 8 is a supervision of the Fig. 7 shown blade and Fig. 9 shows a section of an example not covered by the invention with turbulators and curved cooling passage. DETAILED DESCRIPTION OF THE INVENTION

[0011] Reference will now be made in detail to embodiments of the invention. One or more examples of these embodiments are illustrated in the accompanying drawings. Each of these examples is for illustrative purposes only and is not intended to be limiting of the invention. It will be apparent to those skilled in the art that various changes and variations of the present invention can be made without departing from the scope or spirit of the invention. For example, features shown or described as part of one embodiment may, when taken in conjunction with another embodiment, produce yet another embodiment. It is therefore intended that the present invention cover changes and variations included within the scope of the appended claims and their equivalents.

[0012] The present subject matter is essentially a cooling arrangement for an airfoil of a turbine component. In particular, the present subject matter is a turbine component having one or more curved cooling passages for supplying a coolant to specific regions of the airfoil. In various embodiments, the curved cooling passages may be configured to provide a coolant on or in the vicinity of the airfoil tip of the component. The curved cooling passages may, for example, be formed on the outer periphery of the airfoil (e.g., along the curved shape or profile of the pressure and / or suction side of the airfoil) and in a region substantially adjacent to the airfoil tip.Due to the curvature of the cooling passages and their location on the outer circumference of the blade tip, the coolant flowing through the cooling passages can be effectively directed to the areas on and / or near the blade tip that require cooling. This significantly reduces both the required airflow and the number of cooling openings needed to adequately cool the blade tip.

[0013] The curved cooling passages of the present subject matter are described herein essentially with reference to the rotor blades of a gas turbine. However, those skilled in the art will readily appreciate that the disclosed cooling passages can be formed in essentially any turbine component provided with an airfoil that requires cooling. The curved cooling passages can therefore also be formed, for example, in the rotor blade of a compressor in the compressor section of a gas turbine. Furthermore, the application of the present subject matter is not intended to be limited to gas turbines but also to include steam turbines. It should be further noted that the curved cooling passages can be used both in the components of turbines intended for power generation and in components of turbines used for propulsion in aviation.

[0014] The drawings are explained below. Fig. 1 is a schematic diagram of a gas turbine plant 10. The gas turbine plant 10 essentially comprises a compressor section 12, a plurality of combustion chambers within a combustion chamber section 14, and a turbine section 16. The compressor section 12 and the turbine section 16 may be connected to each other by a shaft 18. During operation of the gas turbine plant 10, compressed air from the compressor section 12 is supplied to the combustion chambers of the combustion chamber section 14. In each combustion chamber, air and fuel are mixed and burned, and the hot combustion gases flow in a hot gas path from the combustion chamber section 14 into the turbine section 16, where energy is extracted from the combustion gases to produce work.

[0015] Fig. 2 shows one embodiment of a turbine blade 20 suitable for use in the turbine section 16 of a gas turbine engine 10. As shown, the blade 20 generally includes a shank portion 22 and an airfoil 24 extending from a substantially planar platform 26. The platform 26 generally serves as the radially inward boundary for the hot combustion gases flowing through the turbine section 16 of the gas turbine engine 10. The shank portion 22 of the blade 20 may be generally configured to extend radially inward from the platform 26 and include sides 28, a cavity 30 defined in part by the sides 28, and one or more angel-wing seals 32 extending horizontally from each side 28.The shaft portion 22 may also include a root structure (not shown), such as a dovetail, for securely attaching the blade 20 to a rotor (not shown) of the turbine portion 16 of the gas turbine system 10.

[0016] The airfoil 24 of the turbine blade 20 may substantially comprise a blade base 34 extending radially outward from the platform 26, and an airfoil tip 36 disposed in the opposite direction to the base 34. The airfoil tip 36 may therefore substantially represent the radially outermost portion of the turbine blade 20. The airfoil 36 may also have a pressure side 38 and a suction side 40 ( Fig. 4) extending between a leading edge 42 and a trailing edge 44. The pressure side 38 may substantially represent an aerodynamic, concave outer surface of the airfoil 24. Similarly, the suction side 40 may substantially represent an aerodynamic, convex outer surface of the airfoil 24.

[0017] Fig. 2 also shows that the turbine blade 20 may also include an airfoil cooling passage 46 extending radially outward from the shank portion 22 toward the blade tip 36 and serving to conduct a coolant—such as air, water, steam, or any other suitable fluid—through the airfoil 24. It should be appreciated that the airfoil cooling passage 46 may be configured in any suitable manner known in the art. For example, in the embodiment shown, the cooling passage 46 includes a plurality of cooling passages 48, 50, 52, 54 extending radially outward from one or more coolant supply passages 56 to a region of the airfoil 24 substantially adjacent the airfoil tip 36.In particular, the airfoil cooling channel 46 includes four radially extending cooling channels 48, 50, 52, 54, which are configured to conduct the coolant supplied from the supply channels 56 throughout the airfoil 24. However, it should be understood by those of ordinary skill in the art that the airfoil cooling channel 46 may include any number of cooling channels 48, 50, 52, 54. It should also be noted that, although the cooling channels 48, 50, 52, 54 are illustrated in the drawing as separate channels extending radially toward the airfoil tip 36, the cooling channels 48, 50, 52, 54 may also be fluidly connected to one another. The airfoil cooling passage 46 may, for example, be configured as a multi-pass cooling passage and may include a plurality of interconnected cooling passages 48, 50, 52, 54 extending radially inward and radially outward within the airfoil 24.In particular, in one embodiment, the cooling channels 48, 50, 52, 54 may form a serpentine path such that the coolant in the channels 48, 50, 52, 54 flows alternately radially outward and radially inward through the airfoil 24. It should also be noted that the airfoil cooling channel 46 may be either an open or a closed cooling circuit. In one embodiment, the airfoil cooling channel 46 may, for example, include a return channel (not shown) that serves to direct the coolant through the shank portion 22 back to the coolant source (not shown). Alternatively, the coolant supplied through the airfoil cooling channel 46 may exit through any number of outlets or openings formed in the airfoil 24.

[0018] Fig. 3 and Fig. 4 illustrate an embodiment of a turbine blade 20 having a plurality of curved cooling passages 60, 62 for cooling a region of the airfoil 24 substantially adjacent the airfoil tip 36 in accordance with aspects of the present subject matter. Fig. 3 shows in particular a partial perspective view of the airfoil 24 of the rotor blade in which a plurality of curved cooling passages 60, 62 are formed. Fig. 4 shows a top view of the Fig. 3, in particular the surface 64 of the blade tip 36.

[0019] In general, the curved cooling passages 60, 62 of the present subject matter may be configured to direct coolant along at least a portion of the outer periphery of the airfoil 24. For example, in various embodiments, the curved cooling passages 60, 62 may be configured to direct coolant along the outer periphery of the airfoil 24 in a region substantially adjacent the airfoil tip 36. Therefore, in the illustrated embodiment, a first curved cooling passage 60 may be formed substantially adjacent the airfoil tip 36 and extend longitudinally along the pressure side 38 of the airfoil 24. Additionally, a second curved cooling passage 62 may be formed substantially adjacent the airfoil tip 36 and extend longitudinally along the suction side 40 of the airfoil 24.However, it should be noted that in alternative embodiments, multiple curved cooling channels may be formed on the pressure side 38 or the suction side 40 and / or extend between the pressure and suction sides 38, 40 of the airfoil 24. Furthermore, in another embodiment, the airfoil 24 may include only a single curved cooling passage formed on the pressure side 38 or the suction side 40 or extending between the pressure and suction sides 38, 40 of the airfoil 24.

[0020] As previously mentioned, the first and second cooling passages 60, 62 may be formed in the airfoil 24 such that coolant may be supplied to portions of the airfoil 24 located at the airfoil tip and / or in the vicinity of the airfoil tip 36. Therefore, in various embodiments, the cooling passages 60, 62 may have any suitable radial position and / or orientation within the airfoil 24 that allows the coolant to be directed toward or substantially adjacent to the substantially horizontal plane defined by the surface 64 of the airfoil tip 36. For example, in the illustrated embodiment, each of the cooling passages 60, 62 may be substantially planar in orientation and extend longitudinally within the airfoil 24 and substantially parallel to the tip surface 64.In such an embodiment, it should be noted that the radial distance 66 of each cooling passage 60, 62 to the tip surface 64 may remain substantially constant between the upstream and downstream ends 68, 70 of the cooling passages 60, 62.

[0021] In another embodiment, each cooling passage 60, 62 may have an angular planar orientation within the airfoil 24 such that the radial distance 66 to the tip surface 64 between the upstream and downstream ends 68, 70 of the cooling passages 60, 62 generally increases or generally decreases. It should also be noted that in alternative embodiments, the cooling passages 60, 62 may be formed in separate planes and / or one or both of the cooling passages 60, 62 may be non-planar in the airfoil 24. Furthermore, it should be noted that the radial distance 66 between the first cooling passage 60 and the tip surface 64 may be equal to or unequal to the radial distance 66 between the second cooling passage 62 and the tip surface 64.

[0022] Furthermore, in various embodiments of the present subject matter, the curvature of the first and second cooling passages 60, 62 may be substantially conformed to the aerodynamic profile of the pressure side 38 and, correspondingly, the suction side 40 of the airfoil 24, such that a coolant may be directed substantially adjacent to the hot outer surfaces of the airfoil 24. For example, in the illustrated embodiment, the first cooling passage 60 may extend longitudinally along a substantial portion of the pressure side 38 of the airfoil 24 and have a varying radius of curvature that substantially conforms to the concave shape or profile of the pressure side 38.Similarly, the second cooling passage 62 may extend longitudinally along a substantial portion of the suction side 40 of the airfoil 24 and may have a varying radius of curvature that substantially conforms to the convex shape or profile of the suction side 40. "Extending longitudinally along a substantial portion of the pressure side 38 and / or the suction side 40 of the airfoil 24" in this context means that the cooling passages 60, 62 extend longitudinally over more than 50% of the width of the pressure side 38 and / or the suction side 40 - defined between the leading edge 42 and the trailing edge 44 of the airfoil 24 (i.e. measured in a direction substantially transverse to the radial direction) - for example over more than 75% of the width between the leading edge 42 and the trailing edge 44 or over more than 90% of the width between the leading edge 42 and the trailing edge 44.The coolant directed through the cooling passages 60, 62 may be delivered to the regions of the airfoil tip 36 that are substantially adjacent to a substantial portion of the outer periphery of the airfoil 24. It should be noted that in other embodiments, the cooling passages 60, 62 need not be configured to extend along a substantial portion of the pressure side 38 and / or the suction side 40. For example, the cooling passages 60, 62 may extend only along a portion of the pressure side 38 and / or the suction side 40 that is less than or equal to 50% of the width between the leading edge 42 and the trailing edge 44. In such embodiments, the cooling passage(s) 60, 62 may generally have a curvature that substantially conforms to the portion of the airfoil 24 along which the cooling passage(s) 60, 62 are formed.It should also be noted that in further embodiments of the present subject matter, the cooling passages 60, 62 may have any suitable curvature and therefore need not have a curvature that substantially conforms to the aerodynamic profile of the airfoil 24.

[0023] As in Fig. 3 and Fig. 4, the first and second cooling passages 60, 62 may be substantially fluidly connected to the cooling channel 46 of the blade so that the coolant flowing through the cooling channel 46 may be directed through the cooling passages 60, 62. Therefore, as shown in Fig. 3 and Fig. 4, one or more transverse channels 72 may be formed in the airfoil 24 between each cooling passage and one or more cooling channels 48, 50, 52, 54 of the airfoil cooling channel 46. Coolant flowing through the airfoil cooling channel 46 may be directed into the cooling passages 60, 62 via the transverse channels 72. It should be noted that in one embodiment, the transverse channels 72 may be formed in the airfoil 24 by drilling a hole from a pressure and / or suction side surface of the airfoil 24 through the cooling passages 60, 62 and into one of the cooling channels 48, 50, 52, 54, with the hole subsequently being sealed on the pressure and / or suction side surface (e.g., by overlay welding).

[0024] Alternatively, the cooling passages 60, 62 may be formed in the airfoil 24 such that the passages 60, 62 are directly fluidly connected to the cooling channels 48, 50, 52, 54 of the cooling channel 46 of the airfoil. Similar to the Fig. 5 and Fig. 6, for example, the upstream ends 68 of the first and second cooling passages 60, 62 may be configured to extend directly into one or more cooling channels 48, 50, 52, 54 of the airfoil cooling channel 46 (e.g., leading edge cooling channel 48) to allow coolant to flow directly from the cooling channel 46 into the cooling passages 60, 62.

[0025] Furthermore, the coolant flowing through each of the cooling passages 60, 62 may exit the passages 60, 62 using substantially any suitable means and / or method known in the art. In one embodiment, the cooling passages 60, 62 may be formed in the airfoil such that the upstream and / or downstream ends 68, 70 of the cooling passages 60, 62 terminate at an outer surface of the airfoil 24. As shown in Fig. 3 and Fig. 4, each cooling passage 60, 62 may have an upstream end 68 having a coolant outlet 74 located substantially adjacent the leading edge 42 of the airfoil 24, and a downstream end 70 having a second coolant outlet 76 common to the cooling passages, located at the trailing edge 44 of the airfoil 24. Consequently, the coolant flowing through the passages 60, 62 may exit through such outlets 74, 76 at the surfaces of the leading edge 42 and the trailing edge 44 of the airfoil 24. However, as already mentioned, the cooling passages 60, 62 need not extend over the entire distance between the leading edge 42 and the trailing edge 44 of the airfoil 24, but can essentially be designed to extend over any portion of the pressure side 38 and the suction side 40 of the airfoil 24.It should therefore be noted that the upstream and downstream ends 68, 70 of the cooling passages 60, 62 may have coolant outlets 74, 76 located substantially anywhere on the outer periphery of the airfoil 24. Furthermore, in further embodiments, one or both cooling passages 60, 62 may have only a single coolant outlet. For example, the cooling passages 60, 62 may be configured such that the upstream or downstream ends 68, 70 of each cooling passage 60, 62 terminate within the airfoil 24.

[0026] In addition or alternatively to the coolant exiting at the ends 68, 70 of the cooling passages 60, 62, the coolant could also exit through numerous film cooling openings 78, 80 formed in the airfoil 24. As shown in Fig. 3 and Fig. 4, for example, numerous film cooling openings 78 may be formed between one or both cooling passages 60, 62 (e.g., the first cooling passage 60) and the pressure side 38 and / or the suction side 40 of the airfoil 24 (e.g., the pressure side 38) to allow the coolant to exit as a film on the pressure and / or suction side surfaces. Alternatively, numerous film cooling openings 80 may be formed in the airfoil between one or both cooling passages 60, 62 (e.g., the second cooling passage 62) and the surface 64 of the airfoil tip 36 to allow the coolant to exit on the tip surface 64. In such embodiments, it should be noted that the coolant flowing through the cooling passages 60, 62 may exit entirely through the film cooling openings 78, 80.For example, the cooling passage(s) 60, 62 in the airfoil may be configured such that the upstream and downstream ends 68, 70 of each cooling passage 60, 62 terminate within the airfoil 24. Alternatively, the outlets 74, 76 formed at the upstream and downstream ends 68, 70 of the cooling passages 60, 62 may simply be closed, for example, by overlay welding or by welding or brazing a plug into the ends 68, 70. It should be noted that the film cooling openings 78, 80 may be oriented in substantially any suitable direction, for example, axially rearward and / or radially upward.

[0027] Fig. 5 and Fig. 6 show an example not encompassed by the invention of a turbine blade 20 with a plurality of curved cooling passages 160, 161, 162 for cooling a region of the blade airfoil 24 that is substantially adjacent to the blade airfoil tip 36. Fig. 5 shows in particular a partial perspective view of the airfoil 24 of the rotor blade with a plurality of curved cooling passages 160, 161, 162 formed therein. Fig. 6 is a supervision of the Fig. 5 and in particular shows the surface 64 of the blade tip 36.

[0028] Similar to the one already mentioned with reference to Fig. 3 and Fig. 4, the curved cooling passages 160, 161, 162 may be formed in the airfoil 24 such that a coolant may be directed to portions of the airfoil 24 located on and / or adjacent to the airfoil tip 36. Therefore, as shown, the cooling passages 160, 161, 162 may be formed in the airfoil 24 substantially adjacent the tip surface 64. For example, the cooling passages 160, 161, 162 may be substantially planar in orientation and extend longitudinally within the airfoil 24 and substantially parallel to the tip surface 64 such that the radial distance 166 from each cooling passage 160, 161, 162 to the tip surface 64 remains substantially constant between the upstream and downstream ends 168, 170 of each cooling passage 160, 161, 162.Alternatively, the cooling channels 160, 161, 162 may be planar and angularly aligned, such that the radial distances 166 increase or decrease as the cooling passages 160, 161, 162 extend through the airfoil 24. In further embodiments, each cooling passage 160, 161, 162 may be arranged in a separate plane and / or some or all of the cooling passages 160, 161, 162 may be non-planarly aligned within the airfoil 24. Furthermore, the radial distance 166 between a particular cooling passage 160, 161, 162 and the tip surface 64 may be equal to or unequal to the radial distances 166 between the other cooling passages 160, 161, 162 and the tip surface 64.

[0029] Furthermore, in various embodiments of the present subject matter, the curvature of the cooling passages 160, 161, 162 may be substantially configured to conform to the aerodynamic profile of the airfoil 24 so that a coolant may be directed along at least a portion of the outer periphery of the airfoil 24. Therefore, in the illustrated embodiment, the cooling passages 160, 161, 162 may be disposed along the pressure side 38 of the airfoil 24 and generally configured to conform to the concave shape or profile of the pressure side 38 of the airfoil 24. For example, each cooling passage 160, 161, 162 may have a different radius of curvature depending on the shape or profile of the portion of the pressure side 38 along which the passage 160, 161, 162 is disposed. As particularly illustrated in Fig. 6, the cooling passage 160 closest to the leading edge 42 may have a relatively small radius of curvature, substantially corresponding to the abrupt curvature of the portion of the pressure side 38 extending directly from the leading edge 42. In contrast, the cooling passage 162 extending toward the trailing edge 40 of the airfoil 24 may have a relatively large radius of curvature, substantially corresponding to the gradual curvature of that portion of the airfoil 24. Accordingly, the coolant conducted through the cooling passages 160, 161, 162 may be supplied to the regions of the airfoil 24 substantially adjacent to the hot outer surface of the pressure side 38 of the airfoil 24.

[0030] In other embodiments, it should be noted that one or more curved cooling passages 160, 161, 162 may also be arranged on the suction side 40 of the airfoil 24 to allow a coolant to be guided along the outer surface of the suction side 40. Furthermore, it should be noted that although three cooling passages 160, 161, 162 are shown in Fig. 5 and Fig. 6, essentially any number of cooling channels may be formed along the pressure side 38 and / or the suction side 40 of the airfoil 24.

[0031] The cooling passages 160, 161, 162 may further be fluidly connected to the airfoil cooling channel 46 to allow coolant to be supplied to the cooling passages 160, 161, 162. As shown, for example, all or a portion of the cooling passages 160, 161, 162 may be arranged in the airfoil 24 such that the upstream ends 168 of the cooling passages 60, 161, 162 are directly fluidly connected to one or more of the cooling channels 48, 50, 52, 54 of the airfoil cooling channel 46. In another embodiment, one or more transverse channels 172 may be arranged in the airfoil 24 between two or more cooling passages 160, 161, 162 such that a portion of the coolant flowing through one of the cooling passages 160, 161, 162 may be redirected to another cooling passage 160, 161, 162. As particularly shown in Fig. 6, a transverse channel 172 may be disposed between the cooling passage 160 closest to the leading edge 38 of the airfoil 24 and the adjacent cooling passage 161 such that a portion of the coolant flowing through the adjacent cooling passage 161 may be diverted into the cooling passage 160 at the leading edge. It should be noted that in one embodiment, the transverse connection 172 may be formed in the airfoil 24 by drilling a hole from the pressure-side surface of the airfoil 24 through the cooling passage 160 and into the adjacent cooling passage 161, with the hole subsequently being sealed on the pressure-side surface (e.g., by overlay welding).

[0032] Alternatively, one or more transverse channels (not shown) may be disposed in the airfoil 24 between the cooling channels 48, 50, 52, 54 of the airfoil cooling channel 46 and each of the cooling passages 160, 161, 162 to allow coolant to be directed from the cooling channel 46 to the cooling passages 160, 161, 162. In still other embodiments, two or more cooling passages 160, 161, 162 may intersect in the airfoil 24 such that coolant can flow directly from one such cooling passage 160, 161, 162 into another.

[0033] Similar to the previously described embodiment, the coolant flowing through each of the cooling passages 160, 161, 162 may exit the cooling passages using any suitable means and / or method known in the art. As shown, the cooling passages 160, 161, 162 may, for example, have outlets 174 for the coolant at their upstream and / or downstream ends 168, 170 so that the coolant may exit at various locations on the outer surface(s) of the airfoil 24. In addition or alternatively to such coolant outlets 174, the coolant may exit through numerous film cooling openings 78, 80 ( Fig. 3 and Fig. 4) disposed in the airfoil 24 between the cooling passages 160, 161, 162 and the pressure side 38 and / or the suction side 40 of the airfoil 24 and / or between the cooling passages 160, 161, 162 and the tip surface 64 of the airfoil 24, so that the coolant can emerge as a film on such outer surfaces.

[0034] Fig. 7 and Fig. 8 show another example not encompassed by the invention of a turbine blade 20 having a plurality of curved cooling passages 260, 261, 262 for cooling a region of the blade airfoil 24 substantially adjacent to the blade airfoil tip 36. Fig. 7 shows a partial side view of the airfoil 24 of the rotor blade with a plurality of curved cooling passages 260, 261, 262 formed therein. Fig. 8 shows a top view of the Fig. 7 and in particular the surface 64 of the blade tip 36.

[0035] As in Fig. 7 and Fig. 8, a plurality of curved cooling passages 260, 261, 262 may be formed on the pressure side 38 of the airfoil 24 and substantially in the vicinity of the airfoil tip 36. Similar to the cooling passages already described with reference to Fig. 5 and Fig. 6, the cooling passages 260, 261, 262 may be configured substantially to conform to the concave shape or profile of the pressure side 38 of the airfoil 24. For example, each cooling passage 260, 261, 262 may have a different radius of curvature depending on the shape or profile of the portion of the pressure side 38 along which the passage 260, 261, 262 is disposed. Accordingly, the coolant conducted through the cooling passages 260, 261, 262 may be supplied to the regions of the airfoil 24 that are substantially adjacent to the hot outer surface of the pressure side 38 of the airfoil 40. In alternative embodiments, it should be noted that one or more cooling passages 260, 261, 262 may also be arranged on the suction side 40 of the airfoil 24 so that a coolant can be guided along the outer surface of the suction side 40.

[0036] Especially in Fig. 7 shows that each of the cooling passages 260, 261, 262 in the airfoil 24 may extend at an angle in the radial direction between its downstream and upstream ends 268, 270. Therefore, the radial distance 266 between the respective cooling passage 260, 261, 262 and the tip surface 64 may vary substantially with the longitudinal extension of the cooling passages 260, 261, 262 through the airfoil 24. For example, in the illustrated embodiment, the cooling passages 260, 261, 262 may extend radially at an upward angle between their downstream and upstream ends 268, 270, such that an outlet 274 for a coolant is formed in the surface 64 of the airfoil tip 36 at the downstream end 270 of each cooling passage 260, 261, 262. The coolant conducted through the cooling passages 260, 261, 262 may exit through the coolant outlets 274 and deposit as a film on the tip surface 64.Alternatively, the cooling passages 260, 261, 262 may extend radially at a downward angle between their downstream and upstream ends 268, 270, such that the coolant outlets 274 are formed at the upstream end 268 of each cooling passage 260, 261, 262. In other embodiments, the upstream and downstream ends 268, 270 of each cooling passage 260, 261, 262 may terminate inside the airfoil 24, such that the coolant is directed substantially adjacent to the tip surface 64 without exiting onto that surface 64. It should also be noted that in various embodiments, each of the cooling passages 260, 261, 262 need not be disposed in the airfoil 24 at an angle relative to the tip surface 64.For example, in a particular embodiment, a portion of the cooling passages 260, 261, 262 may extend longitudinally within the airfoil 24 and substantially parallel to the tip surface 64, while other passages 260, 261, 262 may extend radially upward toward the tip surface 64 or radially downward away from the tip surface 64.

[0037] It should also be noted that, similar to the previously described embodiments, the coolant may be directed to and from the cooling passages 260, 261, 262 using any suitable means and / or method. In the illustrated embodiment, the cooling passages 260, 261, 262 are, for example, directly fluidly connected to the cooling channels 48, 50, 52, 54 of the airfoil cooling channel 44. Alternatively, one or more transverse channels (not shown) may be used to direct the coolant from the airfoil cooling channel 44 to the cooling passages 260, 261, 262 and vice versa, and / or from one of the cooling passages 260, 261, 262 to another. In another embodiment, the cooling passages 260, 261, 262 may be configured to cross each other in the airfoil 24 to allow coolant to flow back and forth between the cooling passages 260, 261, 262.As previously mentioned, the coolant may also exit through the coolant outlets 274 formed in the tip surface 64. Alternatively, the coolant may exit through numerous film cooling openings (not shown) formed in the airfoil 24 between the cooling passages 260, 261, 262 and the pressure side 38 and / or the suction side 40 of the airfoil 24 and / or between the cooling passages 260, 261, 261 and the tip surface 64.

[0038] It should be generally appreciated that the curved cooling passages described herein may be formed in the airfoil 24 using any suitable means known in the art. For example, the curved cooling passages may be formed using an electrical discharge machining (EDM) or casting process. However, in a particular embodiment of the present subject matter, the curved cooling passages are formed using a shaped-tube electrochemical machining (STEM) process. The curved-electrode STEM process is substantially disclosed in patent application Ser. No. 12 / 562528, "Curved Electrode and Electrochemical Machining Method and Assembly Employing the Same," filed September 18, 2009, and assigned to General Electric Company.Unlike conventional STEM drilling, the curved electrode STEM process uses a curved STEM electrode operatively connected to a rotary drive. The rotary drive is configured to move the electrode along a curved path through an object to be machined, for example, through an airfoil 24 of a rotor blade. While the curved electrode rotates along the curved path using the rotary drive, portions of the object to be machined can be removed by spark erosion using a pulsed voltage supplied to the electrode from a voltage source, thereby forming a curved channel in the object.

[0039] It should be appreciated by those of ordinary skill in the art that, in addition to the embodiments described herein, various other cooling patterns and / or arrangements may be used to direct coolant through portions of an airfoil 24 and are also within the scope of the invention. It should therefore be appreciated that the arrangement of the curved cooling passages in the airfoil 24 may be significantly varied to modify the cooling performance of the cooling passages and thus provide preferential cooling of the airfoil 24. Moreover, in various embodiments of the present subject matter, the curved cooling passages in the airfoil 24 may be disposed at locations other than substantially proximate the airfoil tip 36.For example, in a particular embodiment, multiple rows of curved cooling passages may be arranged on the pressure and / or suction sides 38, 40 of the airfoil 24 along the entire radial length of the airfoil 24 between the airfoil base 34 and the airfoil tip 36. An array of curved cooling passages may be provided for the entire airfoil 24.

[0040] It should be further noted that the cooling passages of the present subject matter may have substantially any suitable cross-sectional shape. For example, in the embodiments shown, the cooling passages have a substantially circular cross-section. However, in alternative embodiments, the cooling passages may have an elliptical, flattened, rectangular, or any other suitable non-circular cross-section depending on the desired cooling performance of the cooling passages. The cross-sectional area of ​​each cooling passage may also remain constant or vary along the length of the cooling passage. For example, the size of the tool used to form the cooling channels may be changed during the manufacturing process to change the cross-sectional area of ​​the cooling passages.Therefore, in one embodiment, a curved STEM electrode of a particular size may first be used to form a first portion of a curved cooling passage, and then a smaller curved STEM electrode may be used to form the remainder of the cooling passage.

[0041] In a particular embodiment of the present subject matter, the curved cooling passages may also be provided with turbulators along their length. This means in this context that the surface of the cooling passages may be provided with grooves or ridges or otherwise have a regular surface contour to create turbulence in the flow of the coolant. A sectional view of an embodiment of a cooling passage 360 ​​provided with turbulators is shown in Fig. 9. As shown, the turbulator-equipped cooling passage 360 ​​is equipped with ridges 390 formed along the length of the cooling passage to create turbulence in the flow of coolant. This turbulence can improve the cooling performance of the cooling passage 360 ​​by increasing the heat transfer between the coolant and the blade 24. It should be noted that the ridges 390 are not the Fig.9, but may have substantially any shape and / or arrangement designed to create turbulence in the flow of coolant. Therefore, in alternative embodiments, the ridges 390 may have a substantially square profile and / or may be configured to extend into the airfoil 24 rather than into the cooling passage 360. Additionally, it should be appreciated that the ridges, grooves, or other regular surface contours may be formed in the surface of the cooling passage 360 ​​using any means known in the art. Ridges and / or grooves may be created, for example, by varying the feed rate of the tool used to form the cooling passage 390.Alternatively, in another embodiment, a curved STEM electrode used in the STEM process may be only partially covered with an insulating coating, whereby portions of the electrically conductive part of the electrode come into contact with the surface of the cooling passage 390, causing contouring of the surface.

[0042] Furthermore, it should be noted that the curved cooling passages disclosed herein may also be supplied with coolant from a coolant source other than the airfoil cooling channel 46. For example, the cavity 30 of the shank portion 22 may be pressurized using a coolant—for example, air—to prevent combustion products flowing in the hot gas path from becoming trapped between the turbine blades 20. In such a case, a supply channel (not shown) may be formed through the platform 26 and extend radially upward toward the blade tip 36 of the airfoil 24 to allow the coolant in the cavity 30 to be supplied to the cooling passages.

[0043] A turbine component having a curved cooling passage is disclosed. The turbine component may substantially comprise an airfoil having a base and an oppositely disposed tip. The airfoil may further have a pressure side and a suction side extending between a leading edge and a trailing edge. A cooling passage for the airfoil may be disposed at least partially within the airfoil and configured to conduct a coolant through the airfoil. The curved cooling passage may be substantially fluidly connected to the airfoil cooling passage such that coolant flowing through the airfoil cooling passage may be directed into the cooling passage.The curved cooling passage may also extend substantially longitudinally in the airfoil, between the leading and trailing edges and at least partially along the pressure or suction side of the airfoil. LIST OF REFERENCE SYMBOLS: 10 gas turbines 12 Compressor section 14 Combustion chamber section 16 Turbine section 18 Wave 20 turbine blades 22 Shaft section 24 blades 26 planar platform 28 pages 30 cavity (of the shaft section 22) 32 seals (angel wings) 34 Blade base 36 Blade tip 38 printed pages 40 Suction side 42 leading edge 44 trailing edge 46 Cooling channel of the blade 48 cooling channels 50 cooling channels 52 cooling channels 54 cooling channels 56 coolant supply channels 60 first cooling passage 62 second cooling passage 64 tip surface (of the blade tip 36) 66 radial distance 68 upstream end 70 downstream end 72 cross channels 74 Coolant outlet 76 Coolant outlet 78 film cooling holes 80 film cooling holes 160 Leading edge cooling passage 161 Cooling Passage 162 Cooling Passage 166 radial distance 168 upstream end 170 downstream end 172 cross channels 174 Coolant outlet 260 Leading edge cooling passage 261 Cooling Passage 262 Cooling Passage 266 radial distance 268 upstream end 270 downstream end 360 cooling passage 390 ridges

Claims

[1] Turbine component which includes: an airfoil (24) having a base (34) and a tip (36) disposed in an opposite direction to the base (34), the airfoil (24) further comprising a pressure side (38) and a suction side (40) extending between a leading edge (42) and a trailing edge (44); a cooling channel (46) for the blade (24), which is arranged at least partially in the blade (24) and is adapted to conduct a coolant through the blade (24), and a curved cooling passage (60, 62) formed in the blade (24), wherein the curved cooling passage (60, 62) is fluidly connected to the cooling circuit channel (46) of the blade (24) such that the coolant flowing through the cooling channel (46) of the blade (24) is guided into the curved cooling passage (60, 62), wherein the curved cooling passage (60, 62) extends longitudinally in the airfoil (24) between the leading and trailing edges (42, 44) and at least partially along the pressure side (38) or the suction side (40) of the airfoil (24) between an upstream end (68) and downstream end (70) of the curved cooling passage (60, 62) which terminate at an outer surface of the airfoil (24). [2] The turbine component of claim 1, wherein the curved cooling passage (60, 62) in the airfoil (24) is located substantially in the vicinity of the tip (36). [3] The turbine component of claim 1, wherein the curvature of the curved cooling passage (60, 62) is configured to substantially conform to at least a portion of an aerodynamic profile of the pressure side (38) or the suction side (40) of the airfoil (24). [4] The turbine component of claim 1, wherein the curved cooling passage (60) extends longitudinally along a substantial portion of the pressure side (38) of the airfoil (24), the curved cooling passage (60) having a varying radius of curvature substantially conforming to an aerodynamic profile of the pressure side (38). [5] The turbine component of claim 1, wherein the curved cooling passage (60, 62) is planarly oriented in the airfoil (24) and extends longitudinally in the airfoil (24) substantially parallel to a surface of the tip (36). [6] The turbine component of claim 1, wherein the curved cooling passage (60, 62) extends at an angle in the radial direction between the upstream end (68) and the downstream end (70) of the curved cooling passage (60, 62). [7] The turbine component of claim 1, wherein the curved cooling passage (60, 62) includes a coolant outlet (74, 76) formed in the outer surface of the airfoil (24) such that at least a portion of the coolant conducted through the curved cooling passage (60, 62) exits through the coolant outlet (74, 76). [8] The turbine component of claim 1, wherein a plurality of film cooling openings (78, 80) are formed in the airfoil (24) between the curved cooling passage (60, 62) and the outer surface of the airfoil (24) such that at least a portion of the coolant conducted through the curved cooling passage (60, 62) exits through the plurality of film cooling openings (78, 80). [9] Turbine component comprising: an airfoil (24) having a base (34) and a tip (36) disposed in an opposite direction to the base (34), the airfoil (24) further comprising a pressure side (38) and a suction side (40) extending between a leading edge (42) and a trailing edge (44); a cooling channel (46) for the blade (24), which is arranged at least partially in the blade (24) and is adapted to conduct a coolant through the blade (24), and a plurality of curved cooling passages (60, 62) formed in the blade (24) and substantially in the vicinity of the tip (36), said curved cooling passages (60, 62) being fluidly connected to the cooling circuit channel (46) of the blade (24) such that the coolant flowing through the cooling channel (46) of the blade (24) is directed into said curved cooling passages (60, 62), wherein each curved cooling passage (60, 62) extends longitudinally in the airfoil (24) between the leading and trailing edges (42, 44) and at least partially along the pressure side (38) or the suction side (40) of the airfoil (24) between an upstream end (68) and downstream end (70) of the curved cooling passage (60, 62) terminating at an outer surface of the airfoil (24). [10] The turbine component of claim 9, wherein the curvature of each curved cooling passage (60, 62) is configured to substantially conform to at least a portion of an aerodynamic profile of the pressure side (38) or the suction side (40) of the airfoil (24). [11] The turbine component of claim 9, wherein the curved cooling passages (60) are disposed on the pressure side (38) of the airfoil (24), and wherein each curved cooling passage (60) has a radius of curvature substantially consistent with the portion of the pressure side (38) along which each curved cooling passage (60) is disposed. [12] The turbine component of claim 9, wherein at least one of the curved cooling passages (60, 62) in the airfoil (24) is planarly oriented and extends longitudinally in the airfoil (24) substantially parallel to a surface of the tip (36). [13] Turbine component according to claim 9, wherein at least one of the curved cooling passages (60, 62) extends at an angle in the radial direction between the upstream end (68) and the downstream end (70) of the curved cooling passage (60, 62). [14] The turbine component of claim 9, wherein at least one of the curved cooling passages (60, 62) includes a coolant outlet (74, 76) formed in the outer surface of the airfoil (24) such that at least a portion of the coolant conducted through the curved cooling passage (60, 62) exits through the coolant outlet (74, 76). [15] The turbine component of claim 9, wherein a plurality of film cooling openings (78, 80) are disposed in the airfoil (24) between at least one of the curved cooling passages (60, 62) and the outer surface of the airfoil (24) such that at least a portion of the coolant conducted through the curved cooling passage (60, 62) exits through the plurality of film cooling openings (78, 80).

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