Cooling circuit for turbine components

Abstract

The present invention provides a rotor blade comprising a platform (42) and a shank (36) extending radially inward from the platform (42). The rotor blade also includes a vane (40) extending radially outward from the platform (42). The vane (40) includes a leading edge (52) and a trailing edge (54). A cooling circuit (56) is defined within the shank (36) and the vane (40). The cooling circuit (56) includes a plurality of pins (68). The plurality of pins (68) includes a first set of pins (72) positioned radially inward on the platform (42) and a second set of pins (74) positioned within the vane (40). The cooling circuit (56) also includes a plurality of outlet channels (66) disposed along the trailing edge (54). The plurality of outlet channels (66) are downstream of the plurality of pins (68). The cooling circuit (56) also includes at least one bypass conduit (88) extending from an inlet (90) disposed in the cooling circuit (56) to an outlet (92) positioned on the rear platform surface (116). At least one bypass conduit (88) is positioned radially inside the platform surface (43).

Description

Technical Field

[0001] This disclosure relates in its entirety to cooling circuits for turbine components. More specifically, this disclosure relates to cooling circuits for turbine rotor blades. Background Technology

[0002] Turbines are widely used in fields such as power generation. For example, a conventional gas turbine system includes a compressor section, a combustor section, and at least one turbine section. The compressor section is configured to compress air as it flows through it. The air is then directed from the compressor section to the combustor section, where it is mixed with fuel and burned, producing a hot gas stream. This hot gas stream is supplied to the turbine section, which extracts energy from it to power compressors, generators, and / or other various loads.

[0003] The turbine section typically comprises multiple stages arranged along a hot gas path, such that the hot gas flows through the nozzles and rotor blades of the first stage and through the nozzles and rotor blades of subsequent turbine stages. The turbine rotor blades may be fixed to multiple rotor disks comprising the turbine rotor, wherein each rotor disk is mounted to a rotor shaft to rotate with it.

[0004] Turbine rotor blades typically comprise airfoils extending radially outward from a generally planar platform, and a shank portion extending radially inward from the platform for securing the rotor blades to one of these rotor disks. A cooling circuit circumscribes within the rotor blades to provide a path for cooling air from the compressor section to flow through and cool the high-temperature portions of the airfoils exposed to the hot airflow. In many rotor blades, pin assemblies may be provided within the cooling circuit. Pin assemblies are used to increase convective cooling within the rotor blades by increasing the total surface area exposed to the compressor air. However, using pin assemblies extending radially inward from the platform surface of the rotor blades creates flow path dead zones within the cooling circuit. For example, compressor air may swirl and / or stagnate within the cooling circuit, resulting in unwanted hot spots and degrading overall gas turbine performance. Therefore, rotor blade cooling circuits that allow the use of pin assemblies without causing flow dead zones are desirable in the art. Summary of the Invention

[0005] The aspects and advantages of the components according to this disclosure will be set forth in part in the following description, or may be apparent from the description, or may be learned by practice of the technology.

[0006] According to one embodiment, a turbine component is provided. The turbine component includes a platform and a shank. The platform includes a front platform surface, a rear platform surface, and a platform surface. The shank extends radially inward from the platform. The rotor blades also include vanes extending radially outward from the platform. The vanes include leading edges and trailing edges. A cooling circuit is defined within the shank and the vanes. The cooling circuit includes a plurality of pins. The plurality of pins extend across the cooling circuit. The plurality of pins includes a first set of pins positioned radially inward on the platform surface and a second set of pins positioned within the vanes. The second set of pins is downstream of the first set of pins. The cooling circuit also includes a plurality of outlet channels disposed along the trailing edge. The plurality of outlet channels are downstream of the plurality of pins. The cooling circuit also includes at least one bypass duct extending from an inlet disposed in the cooling circuit to an outlet positioned on the rear platform surface. The at least one bypass duct is positioned radially inward on the platform surface.

[0007] According to another embodiment, a turbine is provided. The turbine includes a compressor section, a combustor section, and a turbine section. A plurality of rotor blades are disposed in the turbine section, and each of the plurality of rotor blades includes a platform and a shank. The platform includes a front platform surface, a rear platform surface, and a platform surface. The shank extends radially inward from the platform. The rotor blades also include vanes extending radially outward from the platform. The vanes include a leading edge and a trailing edge. A cooling circuit is defined within the shank and the vanes. The cooling circuit includes a plurality of pins. The plurality of pins extend across the cooling circuit. The plurality of pins includes a first set of pins positioned radially inward on the platform surface and a second set of pins positioned within the vanes. The second set of pins is downstream of the first set of pins. The cooling circuit also includes a plurality of outlet channels disposed along the trailing edge. The plurality of outlet channels are downstream of the plurality of pins. The cooling circuit also includes at least one bypass duct extending from an inlet disposed in the cooling circuit to an outlet positioned on the rear platform surface. The at least one bypass duct is positioned radially inward on the platform surface.

[0008] These and other features, aspects, and advantages of the components of the present invention will become more readily understood with reference to the following description and the appended claims. The accompanying drawings, which are incorporated in and form a part of this specification, illustrate embodiments of the technology and, together with the description, serve to explain the principles of the technology. Attached Figure Description

[0009] This specification sets forth a complete and practicable disclosure of the components of the invention as understood by one of ordinary skill in the art, with reference to the accompanying drawings, including the best mode for making and using the systems and methods of the invention, wherein:

[0010] Figure 1 This is a schematic diagram of a turbine according to an embodiment of this disclosure;

[0011] Figure 2 A perspective view of a rotor blade according to an embodiment of the present disclosure is shown;

[0012] Figure 3 A cross-sectional top view of a rotor blade according to an embodiment of the present disclosure is shown;

[0013] Figure 4 An enlarged perspective view of a rotor blade according to an embodiment of the present disclosure is shown;

[0014] Figure 5 An enlarged side view of a rotor blade according to an embodiment of the present disclosure is shown; and

[0015] Figure 6 A cross-sectional view of a rotor blade according to an embodiment of the present disclosure is shown. Detailed Implementation

[0016] Reference will now be made in detail to embodiments of the components of the present invention, one or more examples of which are illustrated in the accompanying drawings. Each example is provided by way of explanation of the inventive technique and not as a limitation thereof. Indeed, it will be apparent to those skilled in the art that modifications and variations may be made to the inventive technique without departing from the scope or spirit of the technique protected by the claims. For example, a feature shown or described as part of one embodiment may be used in another embodiment to produce yet another embodiment. Therefore, this disclosure is intended to cover such modifications and variations that fall within the scope of the appended claims and their equivalents.

[0017] The detailed description uses numbers and letters to refer to feature structures in the drawings. Similar or analogous names in the drawings and description have been used to refer to similar or analogous parts of the invention. As used herein, the terms "first," "second," and "third" are used interchangeably to distinguish one part from another and are not intended to indicate the location or importance of the various parts.

[0018] As used herein, the terms “upstream” (or “upward”) and “downstream” (or “downward”) refer to the relative directions of fluid flow within a fluid pathway. For example, “upstream” refers to the direction from which fluid flows, and “downstream” refers to the direction from which fluid flows.

[0019] Sometimes it is necessary to describe parts positioned at different radial locations relative to a central axis. The term "radial" refers to a relative direction that is substantially perpendicular to the axial centerline of a particular part; the term "axial" refers to a relative direction that is substantially parallel and / or coaxially aligned with the axial centerline of a particular part; and the term "circumferential" refers to a relative direction that extends around the axial centerline of a particular part.

[0020] Approximate terms, such as “generally” or “about,” include values ​​that are greater than or less than ten percent of the specified value. When used in the context of angles or directions, such terms include values ​​that are greater than or less than ten degrees of the angle or direction. For example, “generally vertical” includes directions that are within ten degrees of vertical in any direction (e.g., clockwise or counterclockwise).

[0021] Now refer to the attached diagram, Figure 1 A schematic diagram of one embodiment of a turbine is shown, which in the illustrated embodiment is a gas turbine 10. Although industrial or land-based gas turbines are shown and described herein, this disclosure is not limited to industrial and / or land-based gas turbines unless otherwise specified in the claims. For example, turbine components as described herein can be used in any type of turbine, including but not limited to steam turbines, aircraft gas turbines, or marine gas turbines.

[0022] As shown in the figure, the gas turbine 10 generally includes an inlet section 12, a compressor section 14 disposed downstream of the inlet section 12, one or more burners (not shown) disposed in a burner section 16 disposed downstream of the compressor section 14, a turbine section 18 disposed downstream of the burner section 16, and an exhaust section 20 disposed downstream of the turbine section 18. Additionally, the gas turbine 10 may include one or more shafts 22 connecting the compressor section 14 and the turbine section 18.

[0023] The compressor section 14 may generally include a plurality of rotor disks 24 (one of which is shown) and a plurality of rotor blades 26 extending radially outward from each rotor disk 24 and connected to each rotor disk. Each rotor disk 24 may then be coupled to or form part of a shaft 22 extending through the compressor section 14.

[0024] Turbine section 18 may generally include a plurality of rotor disks 28 (one of which is shown) and a plurality of rotor blades 30 extending radially outward from each rotor disk 28 and interconnected to each rotor disk. Each rotor disk 28 may then be coupled to or form a portion of a shaft 22 extending through turbine section 18. Turbine section 18 also includes an outer housing 31 that circumferentially surrounds a portion of shaft 22 and rotor blades 30, thereby at least partially defining a hot gas path 32 through turbine section 18.

[0025] During operation, a working fluid, such as air, flows through inlet section 12 and into compressor section 14, where it is gradually compressed to supply pressurized air to the combustors of compressor section 16. The pressurized air mixes with fuel and burns in each combustor to produce combustion gases 34. Combustion gases 34 flow from combustor section 16 through hot gas path 32 into turbine section 18, where energy (kinetic and / or thermal) is transferred from combustion gases 34 to rotor blades 30, causing shaft 22 to rotate. This mechanical rotational energy can then be used to power and / or generate electricity for compressor section 14. The combustion gases 34 exiting turbine section 18 can then be discharged from gas turbine 10 via exhaust section 20.

[0026] like Figure 2 and Figure 6 As best viewed herein, the gas turbine 10 may define an axial direction A and a circumferential direction C extending about the axial direction A. The gas turbine 10 may also define a radial direction R perpendicular to the axial direction A. As used herein, in some embodiments, the turbine component may be rotor blades 26 and / or 30. In other embodiments, the turbine component may be stator guide vanes (not shown). The function and structure of the stator guide vanes are understood and therefore are not described herein.

[0027] Figure 2 This is a perspective view of an exemplary rotor blade 30 that may be incorporated into one or more embodiments of this disclosure. Figure 2 As shown, the rotor blade 30 typically includes a mounting portion or shank portion 36 having a mounting body 38 and blades 40 extending substantially radially outward from the platform 42. Figures 2 to 6 As shown, platform 42 can be radially positioned between the shank portion 36 and the blade 40. In many embodiments, platform 42 may also include platform surface 43, which can serve as a path 32 for hot gas flowing through turbine section 18. Figure 1 The radial inward boundary of the combustion gas 34.

[0028] In some embodiments, platform surface 43 may be the outermost radial surface of platform 42 and may directly intersect with blade 40. Platform 42 may generally surround blade 40 and may be positioned at the intersection or transition between blade 40 and shank portion 36. Similarly, platform surface 43 may be positioned at the intersection of platform 42 and blade 40. In many embodiments, platform 42 may extend axially beyond shank portion 36.

[0029] Platform 42 may further include a front platform surface 114 facing the combustion gas 34 and a rear platform surface 116 axially separated from the front platform surface 114. The rear platform surface 116 may be downstream of the front platform surface 114. Figure 2As shown, platform 42 may terminate at corresponding front platform surface 114 and rear platform surface 116 in the direction of axis A. The mounting body 38 of the shank portion 36 may extend radially inward from platform 42 and may include root structures (such as dovetails) configured to interconnect or secure rotor blades 30 to rotor disk 28 (e.g., Figure 1 (As shown).

[0030] The winglet 40 may have a generally aerodynamic profile and may include a pressure sidewall 44 and an opposing suction sidewall 46. The curved axis 70 (e.g.) Figure 3 The airfoil 40 (shown) may be defined between the pressure sidewall 44 and the suction sidewall 46, and the curved axis 70 may be generally curved or arcuate. In various embodiments, the pressure sidewall 44 and the suction sidewall 46 may extend substantially radially outward from the root 48 of the airfoil 40 across the platform 42 to the tip 50 of the airfoil 40. The root 48 of the airfoil 40 may be defined at the junction between the airfoil 40 and the platform surface 43. The pressure sidewall 44 generally includes an aerodynamically concave outer surface of the airfoil 40. Similarly, the suction sidewall 46 may generally define an aerodynamically convex outer surface of the airfoil 40.

[0031] The vane 40 may include a leading edge 52 and a trailing edge 54, which are spaced apart from each other and define the end of the vane 40 in the axial direction A. The leading edge 52 of the vane 40 may be the first portion of the vane 40 engaged (i.e. exposed) to combustion gases 34 along the hot gas path 32. The combustion gases 34 may be guided along the aerodynamic profile of the vane 40 (i.e., along the suction sidewall 46 and the pressure sidewall 44) and then discharged at the trailing edge 54.

[0032] The tip 50 is arranged radially opposite to the root 48. Thus, the tip 50 can generally define the radially outermost portion of the rotor blade 30, and can therefore be configured to be positioned adjacent to a stationary shroud or seal (not shown) of the gas turbine 10.

[0033] The handle portion 36 may include a pressure-side impact surface 62 and a suction-side impact surface 64. The pressure-side impact surface 62 may be circumferentially spaced from the suction-side impact surface 64. In some embodiments, the pressure-side impact surface 62 and / or the suction-side impact surface 64 may be generally flat surfaces (which may typically be planar or inclined). In other embodiments, the pressure-side impact surface 62 and / or the suction-side impact surface 64, or at least a portion thereof, may be curved. For example, in Figure 2 In the illustrated embodiment, the pressure-side impact surface 62 or the suction-side impact surface 64 may be bent relative to the axial direction, radial direction, and / or tangential direction.

[0034] The handle portion 36 may further include a leading edge surface 76 axially spaced from a trailing edge surface 78. In some embodiments, the leading edge surface 76 may be positioned within the combustion gas flow 34, and the trailing edge surface 78 may be positioned downstream of the leading edge surface 76. In many embodiments, as shown, the leading edge surface 74 and the trailing edge surface 76 may each be positioned radially inward of the front platform surface 114 and the rear platform surface 116, respectively.

[0035] like Figure 2 As shown, the rotor blades 30 may be at least partially hollow, for example, in the cooling circuit 56 (in Figure 2 (Partially shown in dashed lines) may be enclosed within the vane 40 to guide coolant 58 through the vane 40 between the pressure sidewall 44 and the suction sidewall 46, thereby providing convective cooling thereto. A cooling circuit 56 may be defined within the shank portion 36, the platform 42, and the vane 40, and may include one or more cooling passages 80, 82, 83, 84 for guiding coolant 58 through various portions of the rotor blade 30. For example, the cooling circuit may include one or more leading edge passages 80, one or more intermediate body passages 82, 83, and one or more trailing edge passages 84. Coolant 58 may include coolant from compressor section 14 ( Figure 1 A portion of compressed air and / or steam or any other suitable fluid or gas is used to cool the blades 40. One or more cooling passage inlets 60 are provided along the rotor blades 30. In some embodiments, one or more cooling passage inlets 60 are formed within, along, or by the mounting body 38. The cooling passage inlets 60 are in fluid communication with at least one corresponding cooling passage 80, 82, 83, 84.

[0036] Figure 3 A cross-sectional top view of a rotor blade 30 according to an embodiment of the present disclosure is shown. As shown, the cooling circuit 56 may include a plurality of cooling passages 80, 82, 83, 84 separated by ribs 86. For example, the rotor blade 30 may include one or more leading edge passages 80, one or more intermediate body passages 82, 83 downstream of the leading edge passages 80, and one or more trailing edge passages 84 downstream of the intermediate body passages 82, 83 relative to the direction of the combustion gas flow 34.

[0037] As shown in the figure, the leading edge passage 80 is defined within the rotor blade 30 directly downstream of the leading edge 52 of the vane 40, relative to the direction of flow of the combustion gas 34 on the vane. Similarly, the trailing edge passage 84 is defined within the rotor blade 30 directly upstream of the trailing edge 54 of the vane, relative to the direction of flow of the combustion gas 34 on the vane. The intermediate main passages 82 and 83 are axially defined within the rotor blade 30 between the leading edge passage 80 and the trailing edge passage 84, relative to the curved axis 70.

[0038] like Figure 2 As best shown, coolant 58 can travel substantially radially both inward and outward, passing through cooling circuit 56 and cooling passages 80, 82, 83, 84, to advantageously cool the various cracks, cavities, and sections of rotor blade 30. For example, in Figure 2 In the illustrated embodiment, coolant 58 enters the rotor blade 30 via a cooling passage inlet 60 defined within the mounting body 38 and travels generally radially outward through the intermediate body passage 82 until reaching the tip 50 of the blade 40. At this point, coolant 58 may bend and reverse direction around one or more ribs 86 to continue traveling generally radially inward through another intermediate body air passage 83. Coolant 58 may reverse direction again upon entering the trailing edge passage 84 and travel generally radially outward above the plurality of pins 68 and toward the plurality of outlet passages 66.

[0039] In many implementations, such as Figure 2 In the illustrated embodiment, the vane 40 may define a plurality of outlet channels 66 along its trailing edge 54, said plurality of outlet channels being fluidly coupled to the cooling circuit 56. In some embodiments, the outlet channels 66 may be defined along the trailing edge 54 of the vane 40 and directly fluidly coupled to a trailing edge passage 84. The outlet channels may be spaced apart from each other in the radial direction R and may advantageously provide an outlet for coolant 58 traveling through the cooling circuit 56. The plurality of outlet channels 66 may be formed as substantially hollow cylinders spaced apart from each other and defined between the pressure sidewall 44 and the suction sidewall 46 of the vane 40. Furthermore, as Figure 3 As shown, multiple outlet channels 66 may be oriented along the curved axis 70. The outlet channels 66 may provide an outlet for coolant 58 traveling through the vane 40, to exit the cooling circuit 56. In many embodiments, coolant 58 may be discharged from the outlet channels 66 to mix with combustion gases 34 traveling through the turbine section 18.

[0040] like Figure 2 and Figure 3 As shown, a plurality of pins or pins 68 may be disposed directly upstream of a plurality of outlet channels 66 within the cooling circuit 56 relative to the flow direction of the coolant 58 within the cooling circuit 56. In some embodiments, pins 68 may extend across a trailing edge passage 84. The plurality of pins 68 may extend across the cooling circuit 56 and may be arranged in an array or pattern within the cooling circuit 56. In many embodiments, the plurality of pins 68 may be positioned to allow coolant 58 to pass between and around the pins 68. In some embodiments, the plurality of pins 68 may be used to increase the surface area exposed to convective cooling of the coolant 58 passing through the cooling circuit 56. Each of the plurality of pins 68 may have a substantially circular cross-section. However, in other embodiments (not shown), each pin 68 may have an elliptical, square, rectangular, or any other polygonal cross-sectional shape.

[0041] In some implementations, such as Figures 2 to 5 In the illustrated embodiment, the plurality of pins 68 may include four pin rows 106, 108, 110, and 112, each pin row extending between the shank portion 36 and the tip 50 of the rotor blade 30. For example, the first pin row 106, the second pin row 108, the third pin row 110, and the fourth pin row 112 may be arranged adjacent to each other within the rotor blade 30. Figures 2 to 5 As shown, the first pin row 106 may be the innermost of the four pin rows 106, 108, 110, and 112 in the axial direction. Furthermore, the second pin row 108 may extend axially outward from the first pin row 106, the third pin row 110 may extend axially outward from the second pin row 108, and the fourth pin row 112 may extend axially outward from the third pin row 110. As shown, at least a portion of the fourth pin row 112 may be directly adjacent to the outlet channel 66 within the cooling circuit 56.

[0042] In various embodiments, the plurality of pins 68 may include a first set of pins 72 positioned radially inward of the platform surface 43 (e.g., in the platform 42 and / or the shank portion 36) and a second set of pins 74 disposed in the blade 40. In many embodiments, the second set of pins 74 may be disposed downstream of the first set of pins 72 relative to the direction of coolant 58 flow within the cooling circuit 56 (e.g., generally radially outward). The first set of pins 72 may be disposed within the cooling circuit 56 and between the pressure-side impact surface 62 and the suction-side impact surface 64 of the shank portion 36. In many embodiments, the first set of pins 72 may be disposed radially inward of the platform surface 43 and / or within the shank portion 36 of the rotor blade 30. The second set of pins 74 may be disposed downstream of the first set of pins 72 and upstream of the plurality of outlet channels 66 within the cooling circuit 56. The second set of pins 74 may be disposed radially outward from the first set of pins 72 and the platform surface 43. The second set of pins 74 may extend across the blade 40 ( Figure 3 For example, the second pin assembly 74 may extend between the pressure sidewall 44 and the suction sidewall 46 of the vane 40.

[0043] In many implementations, such as Figure 2 and Figure 3 In the illustrated embodiment, the second pin assembly 74 may be disposed in the trailing edge passage 84 and may extend from the pressure sidewall 44 to the suction sidewall 46, substantially perpendicular to the curved axis 70. Figure 3 As shown, multiple outlet channels 66 can be positioned directly downstream of the second pin group 74 relative to the direction of the combustion gas 34 flowing approximately parallel to the curved axis 70.

[0044] like Figures 2 to 6As shown in the diagram, the rotor blade 30 may further include one or more bypass ducts 88 extending from an inlet 90 disposed within the cooling circuit 56 to an outlet 92 positioned on the rear platform surface 116. The one or more bypass ducts 88 may be formed as hollow cylinders, each hollow cylinder having a trailing edge passage 84 in the cooling circuit 56 intersecting with a hot gas path 32. Figure 1 The bypass conduit 88 provides a passage between the two. The bypass conduit 88 may have a circular cross-sectional shape as shown, or in other embodiments (not shown), the bypass conduit 88 may have an elliptical, square, rectangular or any other polygonal cross-sectional shape.

[0045] One or more bypass ducts 88 may be positioned radially inward of the platform surface 43. In some embodiments, one or more bypass ducts 88 may be defined within both the platform 42 and the shank portion 36 of the rotor blade 30, and may be positioned radially inward of the platform surface 43. In other embodiments, the bypass duct 88 may be completely defined within the shank portion 36 and disposed radially inward from the platform surface 43. In other embodiments, the bypass duct 88 may be completely defined within the platform 42 and disposed radially inward from the platform surface 43.

[0046] like Figures 3 to 6 As shown, at least one bypass conduit 88 can extend from inlet 90 toward suction-side impact surface 64 to outlet 92 disposed on rear platform surface 116. In many embodiments, such as Figure 3 As shown, at least one bypass conduit 88 may extend substantially perpendicular to the suction sidewall 46 of the vane 40.

[0047] like Figure 2 As shown, the inlet 90 of each of the one or more bypass conduits 88 is typically located upstream of the first pin group 72 relative to the flow of coolant 58 within the cooling circuit 56. For example, in some embodiments, the inlet 90 of the bypass conduit 88 may be radially inward from the first pin group 72.

[0048] Each of the one or more bypass conduits 88 may include a constant diameter from inlet 90 to outlet 92. For example, in some embodiments, each of the one or more bypass conduits 88 may have a diameter between about 0.01 inches and about 0.2 inches. In many embodiments, each of the one or more bypass conduits 88 may have a diameter between about 0.025 inches and about 0.175 inches. In other embodiments, each of the one or more bypass conduits 88 may have a diameter between about 0.05 inches and about 0.15 inches. In various embodiments, each of the one or more bypass conduits 88 may have a diameter between about 0.075 inches and about 0.125 inches. In some embodiments, each of the one or more bypass conduits 88 may have a diameter of at most about 0.1 inches.

[0049] In many embodiments, the bypass conduit 88 may be defined within the handle portion 36 and the platform 42, and may extend from an inlet 90 located in the trailing edge passage 84 toward the suction-side impact surface 64 to an outlet 92 disposed on the rear platform surface 116. In certain embodiments, such as Figure 5 As shown, one or more bypass ducts 88 may include a first bypass duct 94 and a second bypass duct 96, each bypass duct having a corresponding inlet 98, 100 within the cooling circuit 56 and a corresponding outlet 102, 104 disposed on the rear platform surface 116. In such embodiments, the corresponding inlet 100 of the second bypass duct 96 may be positioned downstream of the plurality of pins 68 and upstream of the plurality of outlet channels 66 relative to the combustion gas flow 34 above the vane 40. Figure 3 For example, such as Figure 3 As shown, the direction of the combustion gas flow 34 can be from the leading edge 52 to the trailing edge 54 of the vane 40 and is generally parallel to the curved axis 70. In such embodiments, the corresponding inlet 100 of the second bypass duct 96 can be axially positioned relative to the curved axis 70 between the outlet channel 66 and the plurality of pins 68.

[0050] In certain implementation schemes, such as Figures 2 to 5 In the illustrated embodiment, the corresponding inlet 98 of the first bypass conduit 94 can be located directly radially inward from the third pin row 110. Additionally or alternatively, the corresponding inlet 100 of the second bypass conduit 96 can be positioned relative to the distance from the leading edge 52 to the trailing edge 54 of the flap 40. Figure 3 The curved axis 70 (i.e. the direction of combustion gas 34 flow) extends outward from the fourth pin row 112.

[0051] Figure 6A simplified cross-sectional view of the rotor blade 30 according to an embodiment of the present disclosure is shown. As shown, a bypass duct 88 extends from an inlet 90 within a trailing edge passage 84 upstream of the first pin assembly 72 to an outlet 92 disposed on the rear platform surface 116. Furthermore, the bypass duct 88 is completely defined below the platform surface 43 (radially inward), i.e., within the shank portion 36 and the platform 42, and extends generally radially outward toward the platform surface 43 to the outlet 92 positioned on the rear platform surface 116. The bypass duct 88 can advantageously be used to provide a pressure drop within the trailing edge passage 84 that draws at least a portion of the coolant 58 toward itself to achieve a uniform cooling flow distribution.

[0052] In the gas turbine engine 10 ( Figure 1 During operation, cooling fluid flows through the aforementioned passages, cavities, and orifices to cool the rotor blades 30. More specifically, coolant 58 (e.g., bleed air from compressor section 14) enters the rotor blades 30 through the cooling passage inlet 60. Figure 2 The coolant 58 flows through the cooling circuit 56 and various cooling passages 80, 82, 83, 84 to convectively cool both the shank portion 36 and the vane 40 of the rotor blade 30. The cooling fluid 58 flows around and between the pins 68, and then exits the cooling circuit 56 through the outlet passage 66 and / or one or more bypass ducts 88 and flows into the combustion gases 34. Figure 1 Multiple outlet channels 66 are radially outwardly positioned from platform 42 and fluidly connected to cooling circuit 56. Due to the pressure drop generated by the outlet channels 66 within cooling circuit 56, coolant 58 flowing through cooling circuit 56 can travel substantially radially outward and toward outlet channels 66. One or more bypass conduits 88 are used to generate a pressure drop within a portion of cooling circuit 56 defined radially inward from platform surface 43. The pressure drop generated by the one or more bypass conduits 88 advantageously draws at least a portion of coolant 58 over the first pin assembly 72 and toward inlet 90 to form a uniform coolant 58 flow distribution within trailing edge passage 84.

[0053] This written description uses examples to disclose the invention, including the best mode, and also enables any person skilled in the art to practice the invention, including making and using any device or system and performing any combined methods. The patentable scope of the invention is defined by the claims and may include other examples that would occur to a person skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that are not different from the literal language of the claims, or if they include equivalent structural elements that are not substantially different from the literal language of the claims.

Claims

1. A turbine component, comprising: Platform (42), the platform (42) having a front platform surface (114), a rear platform surface (116) and a platform surface (43); Handle (36), the handle extending radially inward from the platform (42); A winglet (40) extends radially outward from the platform (42), the winglet (40) including a leading edge (52) and a trailing edge (54). A cooling circuit (56), defined within the handle (36) and the vane (40), the cooling circuit (56) comprising: Multiple pins (68) extending across the cooling circuit (56) include a first pin group (72) located radially inward on the platform surface (43) and a second pin group (74) located downstream of the first pin group (72) within the winglet (40). Multiple outlet channels (66) are disposed downstream of the multiple pins (68) along the trailing edge (54) of the flap (40); and At least one bypass conduit (88) extends from an inlet (90) disposed in the cooling circuit (56) to an outlet (92) located on the rear platform surface (116), the at least one bypass conduit (88) being located radially inside the platform surface (43).

2. The turbine component according to claim 1, wherein the at least one bypass duct (88) comprises a first bypass duct (94) and a second bypass duct (96), the first bypass duct (94) and the second bypass duct (96) each having a corresponding inlet (98), (100) disposed in the cooling circuit (56) and a corresponding outlet (102), (104) disposed on the rear platform surface (116).

3. The turbine component according to claim 2, wherein, relative to the curved line (70) from the leading edge (52) to the trailing edge (54) of the blade (40), the corresponding inlet (100) of the second bypass duct (96) is downstream of the plurality of pins (68) and upstream of the plurality of outlet channels (66).

4. The turbine component according to claim 1, wherein the cooling circuit (56) includes a leading edge passage (80), a middle body passage (82), (83) and a trailing edge passage (84), wherein the inlet (90) of the at least one bypass duct (88) is disposed in the trailing edge passage (84).

5. The turbine component according to claim 1, wherein the handle (36) includes a suction-side impact surface (64) circumferentially separated from the pressure-side impact surface (62).

6. The turbine component according to claim 5, wherein the at least one bypass duct (88) extends from the inlet (90) toward the suction-side impact surface (64) to the outlet (92).

7. The turbine component according to claim 1, wherein the at least one bypass duct (88) is defined within the platform (42) and the handle (36).

8. The turbine component of claim 7, wherein the at least one bypass duct (88) has a diameter between about 0.01 inches and about 0.2 inches.

9. The turbine component according to claim 1, wherein the pins of the second pin group (74) of the plurality of pins (68) extend from the suction sidewall (46) to the pressure sidewall (44).

10. The turbine component according to claim 1, wherein the plurality of pins (68) are arranged in four rows extending between the shank (36) and the top of the blade (40).

11. A turbine, comprising: Compressor section; Burner section; Turbine section; Multiple rotor blades, wherein the multiple rotor blades are disposed in the turbine section, each of the multiple rotor blades comprising: Platform (42), the platform (42) having a front platform surface (114), a rear platform surface (116) and a platform surface (43); Handle (36), the handle extending radially inward from the platform (42); A winglet (40) extends radially outward from the platform (42), the winglet (40) including a leading edge (52) and a trailing edge (54). A cooling circuit (56), defined within the handle (36) and the vane (40), the cooling circuit (56) comprising: Multiple pins (68) extending across the cooling circuit (56) include a first pin group (72) located radially inward on the platform surface (43) and a second pin group (74) located downstream of the first pin group (72) within the winglet (40). Multiple outlet channels (66) are disposed downstream of the multiple pins (68) along the trailing edge (54) of the flap (40); and At least one bypass conduit (88) extends from an inlet (90) disposed in the cooling circuit (56) to an outlet (92) located on the rear platform surface (116), the at least one bypass conduit (88) being located radially inside the platform surface (43).

12. The turbine according to claim 11, wherein the at least one bypass duct (88) comprises a first bypass duct (94) and a second bypass duct (96), the first bypass duct (94) and the second bypass duct (96) each having a corresponding inlet (98), (100) disposed in the cooling circuit (56) and a corresponding outlet (102), (104) disposed on the rear platform surface (116).

13. The turbine according to claim 12, wherein, relative to the curved line (70) from the leading edge (52) to the trailing edge (54) of the blade (40), the corresponding inlet (100) of the second bypass duct (96) is downstream of the plurality of pins (68) and upstream of the plurality of outlet channels (66).

14. The turbine according to claim 11, wherein the cooling circuit (56) comprises a leading edge passage (80), a middle body passage (82), (83) and a trailing edge passage (84), wherein the inlet (90) of the at least one bypass duct (88) is disposed in the trailing edge passage (84).

15. The turbine of claim 11, wherein the shank (36) includes a suction-side impact surface (64) circumferentially separated from the pressure-side impact surface (62).

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