Turbomachine rotor blade with cooling circuit featuring offset ribs

Abstract

To provide a rotor blade for a turbomachine.SOLUTION: A rotor blade (328) for a turbomachine (10) is provided. The rotor blade (328) includes a platform (342), an airfoil (340) extending from the platform (342), and a cooling circuit (372) extending within the platform (342) and the airfoil (340). The cooling circuit (372) includes a plurality of cooling passages (356) defined by a plurality of ribs (374). The ribs (374) include an offset rib (380).SELECTED DRAWING: Figure 3

Description

Technical Field

[0001] The present disclosure generally relates to turbomachines, and more particularly to airfoils for turbomachines that include a cooling circuit having offset ribs.

Background Art

[0002] Turbomachines are used in various industries and applications for energy transfer purposes. For example, a gas turbine engine generally includes a compressor section, a combustion section, a turbine section, and an exhaust section. The compressor section gradually increases the pressure of the working fluid entering the gas turbine engine and supplies this compressed working fluid to the combustion section. The compressed working fluid and fuel (e.g., natural gas) are mixed within the combustion section and burned within the combustion chamber to generate high-pressure and high-temperature combustion gases. The combustion gases flow from the combustion section to the turbine section, where they expand to generate work. For example, the expansion of the combustion gases in the turbine section can rotate a rotor shaft connected to a generator to generate electricity. The combustion gases are then discharged from the gas turbine through the exhaust section.

[0003] The turbine section typically includes rows of circumferentially spaced stator vanes and rotor blades. The rotor blades generally have a pressure side and a suction side and include airfoils that extend radially upward from a platform. The shank portion can extend radially downward from the platform and can include a structure for fixing the rotor blade to the turbine wheel. The platform generally defines an inner boundary of the high-temperature combustion gases flowing through the gas path. The intersection of the platform and the airfoil can be an area of high stress concentration due to high-temperature combustion gases, mechanical loads thereon, and other causes.

[0004] More specifically, significant thermal or otherwise induced strain is often present at the intersection of the airfoil and platform. This induced strain can result from temperature differences between the airfoil and platform, temperature differences between the positive and negative pressure sides, and rotational speed loads. Combined with geometric discontinuities within the region, this induced strain can create areas of extremely high stress that can limit the overall component life. To date, these problems have been addressed by attempting to keep geometric discontinuities such as root turns, tip turns, and internal ribs away from the intersection. Furthermore, temperature control around the intersection has been attempted. However, the airfoil-platform intersection often continues to limit the life of many rotor blades.

[0005] Therefore, improved features for cooling and / or reducing distortion of rotor blades are desired in the art. [Overview of the project]

[0006] The aspects and advantages of the assemblies provided in this disclosure are partially described in the following description, or become apparent from the description, or can be learned through the practice of the art.

[0007] According to one embodiment, a rotor blade for a turbomachinery is provided. The rotor blade includes a platform, an airfoil extending from the platform, and a cooling circuit extending within the platform and the airfoil. The cooling circuit includes a plurality of cooling passages defined by a plurality of ribs. The plurality of ribs include offset ribs.

[0008] According to another embodiment, a turbomachinery is provided. The turbomachinery defines an axial direction, a circumferential direction extending around the axial direction, and a radial direction perpendicular to the axial direction. The turbomachinery includes a compressor, a combustor downstream of the compressor, and a turbine downstream of the combustor. The turbine includes rotor blades mounted on a rotor disk. The rotor blades include a platform, an airfoil extending radially outward from the platform, and a cooling circuit extending within the platform and the airfoil. The cooling circuit includes a plurality of cooling passages defined by a plurality of ribs. The plurality of ribs include offset ribs.

[0009] These and other features, aspects, and advantages of this assembly will be better understood by referring to the following description and the appended claims. The appended drawings, which are incorporated herein and constitute part of this specification, illustrate embodiments of the art and, together with the description in the specification, help to illustrate the principles of the art.

[0010] A complete and implementable disclosure of this assembly, including best modes of fabrication and use of the system and method intended for those skilled in the art, is described herein with reference to the accompanying drawings. [Brief explanation of the drawing]

[0011] [Figure 1] This is a schematic diagram of a turbomachinery according to an embodiment of the present disclosure. [Figure 2] This is a side view of an exemplary rotor blade according to an embodiment of the present disclosure. [Figure 3] This is an exemplary partially cutaway perspective view of a rotor blade according to an embodiment of the present disclosure. [Figure 4] This is a plan view of an exemplary rotor blade as seen radially inward, according to an embodiment of the present disclosure. [Figure 5] This is an enlarged cross-sectional view of a portion of the rotor blade along line 5-5 in Figure 4. [Figure 6]This is a radially inward cross-sectional view of an exemplary rotor blade along line 6-6 in Figure 2, according to an embodiment of the present disclosure. [Figure 7] This is a radially inward cross-sectional view of an exemplary rotor blade platform along line 7-7 in Figure 2, according to an embodiment of the present disclosure. [Figure 8] Figures 4 and 5 show enlarged cross-sectional views of a portion of the rotor blade along line 8-8. [Modes for carrying out the invention]

[0012] Hereinafter, embodiments of the assembly are given in detail, one or more examples of which are shown in the drawings. Each example is provided for illustrative purposes of the Art and is not intended to limit the Art. Indeed, it will be apparent to those skilled in the art that modifications and changes are possible in the Art without departing from the scope or spirit of the claimed Art. For example, features illustrated or described as part of one embodiment can also be used in another embodiment to bring about further embodiments. Thus, this disclosure is intended to encompass such modifications and changes within the scope of the appended claims and their equivalents.

[0013] Detailed descriptions use numerals and letters to refer to features in the drawings. Similar or identical reference numerals in the drawings and descriptions are used to refer to similar or identical parts of the present invention. As used herein, the terms “first,” “second,” and “third” may be used interchangeably to distinguish one component from another and are not intended to indicate the location or importance of any individual component.

[0014] As used herein, the terms “upstream” (or “forward”) and “downstream” (or “backward”) refer to relative directions of fluid flow in a fluid path. For example, “upstream” refers to the direction in which the fluid is flowing, and “downstream” refers to the direction in which the fluid is flowing. The term “radially” refers to a relative direction substantially perpendicular to the axial centerline of a particular component; the term “axially” refers to a relative direction substantially parallel and / or coaxial with the axial centerline of a particular component; and the term “circumferentially” refers to a relative direction extending around the axial centerline of a particular component. Approximate terms such as “generally” or “about” include values ​​within plus or minus 10 percent of the stated value. When used in the context of angles or directions, such terms include a range of plus or minus 10 degrees of the stated angle or direction. For example, “generally perpendicular” includes any direction, e.g., within 10 degrees from perpendicular in a clockwise or counterclockwise direction.

[0015] Referring here to the drawings, Figure 1 shows a schematic diagram of one embodiment of a turbomachinery, which in the illustrated embodiment is a gas turbine 10. Although industrial and onshore gas turbines are shown and described herein, this disclosure is not limited to onshore and / or industrial gas turbines unless specifically stated in the claims. For example, the rotor blades described herein can be used in any type of turbomachinery, including, but not limited to, steam turbines, aircraft gas turbines, or marine gas turbines.

[0016] As shown, the gas turbine 10 generally includes an inlet section 12, a compressor section 14 located downstream of the inlet section 12, a number of combustors (not shown) in a combustor section 16 located downstream of the compressor section 14, a turbine section 18 located downstream of the combustor section 16, and an exhaust section 20 located downstream of the turbine section 18. In addition, the gas turbine 10 may include one or more shafts 22 coupled between the compressor section 14 and the turbine section 18.

[0017] 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 24. Each rotor disk 24 may then be coupled to or form part of a shaft 22 extending through the compressor section 14.

[0018] The 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 connected to each rotor disk 28. Each rotor disk 28 may then be coupled to or form part of a shaft 22 extending through the turbine section 18. The turbine section 18 further includes an outer casing 31 that circumferentially surrounds part of the shaft 22 and the rotor blades 30, thereby at least partially defining a hot gas path 32 through the turbine section 18.

[0019] During operation, a working fluid such as air flows through the inlet section 12 into the compressor section 14, where the air is gradually compressed, thereby supplying pressurized air to the combustors in the combustor section 16. The pressurized air is mixed with fuel and burned in each combustor to produce combustion gases 34. The combustion gases 34 flow from the combustor section 16 to the turbine section 18 through the high-temperature gas path 32, where energy (kinetic and / or thermal energy) is transferred from the combustion gases 34 to the rotor blades 30, causing the shaft 22 to rotate. This mechanical rotational energy can then be used to power the compressor section 14 and / or generate electricity. The combustion gases 34 discharged from the turbine section 18 can then be exhausted from the gas turbine 10 through the exhaust section 20.

[0020] As can be seen in FIGS. 2 and 3, the turbomachine 10 can define an axial direction A and a circumferential direction C that extends around the axial direction A. The turbomachine 10 can also define a radial direction R that is perpendicular to the axial direction A.

[0021] FIG. 2 is a side view of an exemplary rotor blade 328 according to one or more embodiments of the present disclosure. FIG. 3 is a perspective view of the exemplary rotor blade 328 of FIG. 2. As shown in FIGS. 2 and 3, the rotor blade 328 generally includes a mounting or shank portion 336 having a mounting body 338 and an airfoil portion 340 that extends outwardly from a substantially flat platform 342, e.g., generally along the radial direction R. The platform 342 generally functions as a radially inner boundary for the hot combustion gas 34 that flows through the hot gas path 32 of the turbine section 18 (FIG. 1). The platform 342 extends along the axial direction A from a front face 320 to a rear face 322. As shown in FIG. 3, the mounting body 338 of the mounting or shank portion 336 can extend radially inward from the platform 342 and can include a root structure, such as a double-tail, configured to interconnect or secure the rotor blade 328 to the rotor disk 28 (FIG. 1).

[0022] The airfoil section 340 includes a positive pressure sidewall 344 and an opposing negative pressure sidewall 346. The positive pressure sidewall 344 and the negative pressure sidewall 346 extend substantially radially outward from the platform 342, from the root 348 of the airfoil section 340 to the tip 350 of the airfoil section 340, which can be defined at the intersection between the airfoil section 340 and the platform 342. The positive pressure sidewall 344 connects to the negative pressure sidewall 346 at the leading edge 352 and the trailing edge 354 downstream of the leading edge 352, so that the airfoil section 340 extends between the leading edge 352 and the trailing edge 354. The positive pressure sidewall 344 generally provides an aerodynamically concave outer surface of the airfoil section 340. Similarly, the negative pressure sidewall 346 can generally define an aerodynamically convex outer surface of the airfoil section 340. The tip 350 is positioned radially opposite the root 348. Therefore, the tip 350 can generally define the radially outermost portion of the rotor blade 328, and can be configured to be positioned adjacent to a stationary shroud or seal (not shown) of the turbomachinery 10. The tip 350 may include a tip cavity 366.

[0023] As shown in Figure 3, the rotor blade 328 may be at least partially hollow, and for example, the rotor blade 328 may include a defined cooling circuit 372 inside. The cooling circuit 372 is circumscribable within the rotor blade 328 and may include a plurality of cooling passages 356 (partially dashed in Figure 3) that can deliver coolant 358 through the airfoil portion 340 between the positive pressure sidewall 344 and the negative pressure sidewall 346, thereby providing convective cooling. The cooling passages 356 may be defined at least partially by a plurality of ribs 374, and between them. The ribs 374 may extend generally along the radial direction R, partially through the cooling circuit 372, as shown in Figure 3. The ribs 374 may extend entirely through the cooling circuit 372 between the positive pressure sidewall 344 and the negative pressure sidewall 346, as shown in Figure 6. Thus, the plurality of ribs 374 can partition the cooling circuit 372 and at least partially form or define the cooling passages 356. For example, each rib 374 can be terminated radially near either the root turn 376 or the tip turn 378.

[0024] As shown in FIG. 3, the cooling circuit 372 may be a serpentine cooling circuit including a plurality of root turn portions 376, such as approximately 180° turn portions proximate to the root 348 of the airfoil portion 340, and a plurality of tip turn portions 378, such as approximately 180° turn portions proximate to the tip 350 of the airfoil portion 340. The serpentine cooling circuit 372 extends along a first cooling passage 356 of the plurality of cooling passages 356 in a first direction, extends in the opposite direction through the tip turn portion 378, extends along a second cooling passage 356 of the plurality of cooling passages 356 in a second direction generally opposite to the first direction, extends in the opposite direction through the root turn portion 376 again, returns in the first direction along a third cooling passage 356 of the plurality of cooling passages 356, and may be repeated any number of times in a similar manner hereinafter.

[0025] The coolant 358 may include a portion of the compressed air from the compressor section 14 (FIG. 1) and / or vapor or any other suitable gas or other fluid for cooling the airfoil portion 340. One or more coolant inlets 360 are disposed along the rotor blade 328. In some embodiments, one or more coolant inlets 360 are formed within, along, or by the mounting body 338. The coolant inlet 360 is in fluid communication with at least one corresponding cooling passage 356. A plurality of coolant outlets 364 may be in fluid communication with the tip cavity 366. Each cooling passage 356 is in fluid communication with at least one of the coolant outlets 364. In some embodiments, the tip cavity 366 may be at least partially surrounded by a positive pressure side tip rail 368 and a negative pressure side tip rail 370.

[0026] As seen in FIG. 3, the cooling passages 356 extend into each of the shank portion 336 and the airfoil portion 340. For example, the cooling passage 356 may extend from the shank portion 336 to the airfoil portion 340, such as from one or more coolant inlets 360 in the shank portion 336 to at least one coolant outlet 364 at the tip 350 of the airfoil portion 340, between the shank portion 336 and the airfoil portion 340.

[0027] The rotor blade 328 may also include an offset rib 380. The offset rib 380 may be positioned close to the leading edge 352 of the airfoil section 340, for example, in the foremost or turned portion of the cooling circuit 372, closer to the leading edge 352 than the trailing edge 354. For example, as shown in Figure 5, the offset rib 380 may be offset along the axial direction A. For example, the offset rib 380 may include an outer portion 384 that is generally aligned radially R and / or with a plurality of ribs 374, and an offset portion 385 that is radially inward of the outer portion 384 (for example, closer to the root 348 of the airfoil section 340 and / or the shank portion 336 of the rotor blade 328). The offset portion 385 of the offset rib 380 may be offset by a distance 388 along the axial direction A, as shown in Figure 5. The offset rib 380 may be offset rearward, for example, downstream, toward the trailing edge 354 of the airfoil section 340 and / or toward the rear surface 322 of the platform 342. The offset rib 380 may extend inward along a radial R to a tip 386 that defines the inner end of the rib 380. The portion of the cooling circuit 372 inside the tip 386 of the offset rib 380 may be continuous from one cooling passage 356 to the next adjacent cooling passage 356, thereby defining a turn in the cooling circuit 372. More specifically, if the tip 386 of the offset rib 380 is close to the root 348 of the airfoil section 340, for example, closer to the root 348 than to the tip 350, the turn in the cooling circuit 372 defined by the offset rib 380 may be a root turn 376.

[0028] Figure 4 is a plan view of an exemplary rotor blade 328 viewed radially inward. Figure 5 is an enlarged longitudinal cross-sectional view of a portion of the rotor blade 328 along line 5-5 in Figure 4. Figure 8 is an enlarged cross-sectional view of a portion of the rotor blade 328 along the cross-sectional plane indicated by line 8-8 in Figures 4 and 5. For example, as seen in Figures 5 and 8, the offset rib 380 generally separates and defines two adjacent cooling passages 356 of a plurality of cooling passages 356. In some embodiments, for example, as shown in Figures 5 and 8, the offset rib 380 may also include one or more perforations 382 extending through the offset rib 380 from one side to the other, thereby providing, for example, one or at least one additional passage for the coolant 358 through one cooling passage 356 to the next cooling passage 356, in addition to the root turn 376. In embodiments where a plurality of perforations 382 are provided, the size of the perforations 382 may vary. For example, the proportion of the perforations 382 may decrease as the width or thickness 392 of the rib 380 decreases.

[0029] Figure 6 is a cross-sectional view of the rotor blade 328, viewed radially inward, passing through the airfoil 340 along line 6-6 in Figure 2. Figure 7 is a cross-sectional view of the rotor blade 328, viewed radially inward, passing through the platform 342 along line 7-7 in Figure 2. For example, as seen in Figures 5-7, the offset rib 380 may define a thickness 392, for example, generally along the axial direction A and / or along the flow direction defined from the leading edge 352 to the trailing edge 354 of the airfoil 340. In some embodiments, for example, as shown in Figures 5-7, the thickness 392 of the offset rib 380 may increase as it moves inward along the radial direction R, for example, towards the root 348 of the airfoil 340 and / or towards the shank portion 336 of the rotor blade 328.

[0030] For example, as shown in Figures 6 and 7, the airfoil 340 may define a camber line 390 extending from the leading edge 352 to the trailing edge 354. Those skilled in the art will recognize that the term “camber line” refers to the midpoint or mean line of the airfoil profile, for example, the midpoint between the positive-pressure side surface and the negative-pressure side surface. In this disclosure, the positive-pressure side surface and the negative-pressure side surface of the airfoil 340 are defined by a positive-pressure sidewall 344 and a negative-pressure sidewall 346, respectively.

[0031] The offset rib 380 can define an angle θ ("theta") with respect to the camber line 390. In some embodiments, the angle θ may be generally perpendicular to the camber line 390 near the tip 350 of the airfoil section 340, or it may vary along the radial direction R. For example, the angle θ may increase as the offset rib 380 moves radially inward (e.g., towards the root 348 of the airfoil section 340 and / or the shank portion 336 of the rotor blade 328). For example, the offset rib 380 may converge with or branch off from the camber line 390, which moves radially inward toward the lower shank 336, such that the angle θ increases or decreases inward along the radial direction R. Therefore, the dimensions of the offset rib 380 extending between the positive pressure sidewall 344 and the negative pressure sidewall 346 (for example, the width 394 of the offset rib 380, as described below) can be increased, thereby increasing the structural compliance of the offset rib 380 with respect to the rotor blade 328, for example, its airfoil portion 340.

[0032] As shown in Figure 8, the cooling circuit 372, for example, the cooling passage 356 and / or its root turn 376, can define a width 394. The width 394 of the cooling circuit 372 may be defined along the circumferential direction C, along a direction generally perpendicular to the camber line 390, and / or along the direction between the positive pressure sidewall 344 and the negative pressure sidewall 346 (for example, generally perpendicular to the positive pressure sidewall 344 and the negative pressure sidewall 346). As seen in Figure 8, in some embodiments, the width 394 of the cooling circuit 372, for example, the cooling passage 356 partially defined by the offset rib 380, may increase inward along the radial direction R, for example toward the root 348 of the airfoil section 340, and / or toward the shank section 336 of the rotor blade 328.

[0033] As can be seen by comparing Figure 5 with Figure 8, the offset rib 380 is shown in the background of Figure 8. Furthermore, the offset rib 380 is angled so as to be away from the cross-sectional plane (within the page) of Figure 8. Therefore, it should be understood that the cooling passage 356 shown in Figure 8 is shown in the foreground, for example, in front of the offset rib 380 that partially defines the cooling passage 356, so that the arrow for 356 in Figure 8 points to the space in front of the offset rib 380.

[0034] In some embodiments, the root turn portion 376 may be rectangular, for example, oval in shape. For example, as seen in Figure 8, the tip 386 of the offset rib 380 defining the radially outer boundary of the root turn portion 376 may be inclined or asymmetrical. Thus, for example, the end portion of the offset rib 380, for example, the tip 386, can define a non-circular arcuate boundary of the root turn portion 376. Such a geometric shape can provide a smooth stress flow contour.

[0035] This specification discloses the present invention, including its best mode, and uses examples to enable any person skilled in the art to practice the invention, including the fabrication and use of any device or system and the implementation of any incorporated method. The patentable scope of the present invention is defined by the claims and may include other embodiments that a person skilled in the art may conceive. Such other embodiments are intended to be within the claims if they include structural elements that are not different from the language of the claims, or equivalent structural elements that do not substantially differ from the language of the claims. [Explanation of Symbols]

[0036] 10 Gas Turbines / Turbo Machinery 12 Entrance Section 14 Compressor Section 16. Combustor Section 18 Turbine Section 20 Exhaust Section 22 shafts 24 Rotor Discs 26 rotor blades 28 Rotor Discs 30 rotor blades 31 Outer casing 32 High-temperature gas pathway 34 Combustion gases 320 front 322 Rear 328 Rotor Blades 336 Shank section / Lower shank 338 Mounting Unit 340 Airfoil / airfoil section 342 platforms 344 Positive pressure sidewall 346 Negative pressure sidewall 348 Root 350 Tip 352 Leading edge 354 Trailing edge 356 Cooling passage / arrow 358 Coolant 360 Cooling passage entrance 364 Coolant outlet 366 Tip Cavity 368 Positive pressure side end rail 370 Negative pressure side end rail 372 Cooling Circuit / Serpentine Cooling Circuit 374 Rib 376 Root turn section 378 Tip Turn Section 380 Offset Rib 382 Perforation 384 Outer part 385 Offset portion 386 Tip 388 distance 390 Camber line 392 Thickness 394 width 5-5 line 6-6 line 7-7 line 8-8 line A-axis C Circumferential direction R Radial direction θ angle

Claims

1. A rotor blade (328) for a turbomachine (10), the rotor blade (328) comprising: a platform (342), an airfoil (340) extending from the root (348) to the tip (350) of the platform (342), a cooling circuit (372) extending into the platform (342) and the airfoil (340), the cooling circuit (372) comprising a plurality of cooling passages (356) defined by a plurality of ribs (374), the plurality of ribs (374) comprising offset ribs (380); and the offset ribs (380) define and separate two adjacent cooling passages of the plurality of cooling passages (356), the offset ribs (380) include an outer portion (384) and an offset portion (385) radially inside the outer portion (384), the outer portion (384) being aligned with the radial direction of the turbomachine (10) and extending between the tip (350) of the airfoil (340) and the offset portion (385), the offset portion (385) being offset along the axial direction (A) towards the trailing edge (354) of the airfoil (340) and extending radially inward along the radial direction to a tip (386) defining the inner end of the offset rib (380), a portion of the cooling passage inside the tip (386) of the offset rib (380) between the adjacent cooling passages is continuous, and the offset rib (380) defines a root turn portion (376) of the cooling circuit (372), a perforation (382) passing through the offset rib (380) is defined in the offset rib (380), the rotor blade (328).

2. The rotor blade (328) according to claim 1, wherein the perforation (382) is defined in an offset portion (385) of the offset rib (380) radially outside the root turn portion (376).

3. The rotor blade (328) according to claim 1, wherein the thickness (392) of the offset rib (380) increases toward the root (348) of the airfoil portion (340).

4. The rotor blade (328) according to claim 1, wherein the end portion of the offset rib (380) defines a non-circular arc-shaped boundary of the root turn portion (376) of the cooling circuit (372).

5. The rotor blade (328) according to claim 1, wherein the offset rib (380) is positioned in close proximity to the leading edge (352) of the airfoil portion (340).

6. The rotor blade (328) according to claim 1, wherein the width (394) of the cooling circuit (372) of the offset rib (380) increases toward the root (348) of the airfoil portion (340).

7. The rotor blade (328) according to claim 1, wherein the airfoil portion (340) extends from a leading edge (352) to a trailing edge (354), defines a camber line (390) passing through the leading edge (352) and the trailing edge (354), and the offset rib (380) is positioned at an angle (θ) with respect to the camber line (390) of the airfoil portion (340), and the angle (θ) changes radially inward.

8. The rotor blade (328) according to claim 1, wherein the cooling circuit (372) comprises a serpentine cooling circuit (372).

9. A turbomachinery (10) defining an axial direction (A), a circumferential direction (C) extending around the axial direction (A), and a radial direction (R) perpendicular to the axial direction (A), Compressor and, The combustor downstream of the compressor, A turbine downstream of the combustor, wherein the turbine comprises a rotor blade (328) according to any one of claims 1 to 8 attached to a rotor disk (24) A turbomachinery (10) equipped with the above.

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