Damper stack for turbomachinery rotor blades - Patents.com

Abstract

To provide damper stacks for use internally in rotor blades.SOLUTION: Damper stacks (70), rotor blades (30) and turbomachines are provided. A rotor blade includes a main body including a shank (38) and an airfoil (36) that extends radially outwardly from the shank (38). The rotor blade (30) further includes a platform (42) surrounding the main body, the platform (42) comprising a slash face (56, 58). The rotor blade (30) further includes a damper stack (70) disposed at the slash face (56, 58) and extending generally along an axial direction. The damper stack (70) includes a plurality of damper pins (72), each of the damper pins (72) being in contact with a neighboring damper pin (72).SELECTED DRAWING: Figure 2

Description

[Technical field]

[0001] The present disclosure relates generally to rotor blades for turbomachines, and more particularly, to damper stacks for use within rotor blades. [Background technology]

[0002] Turbomachines are utilized in various industries and applications for the purpose of energy transfer. 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 a working fluid entering the gas turbine engine and supplies the compressed working fluid to the combustion section. The compressed working fluid and fuel (e.g., natural gas) are mixed in the combustion section and combusted in a 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, for example, a generator to generate electricity. The combustion gases are then exhausted from the gas turbine via the exhaust section.

[0003] The compressor and turbine sections generally include multiple rotor blades, typically arranged in multiple stages. During engine operation, vibrations may be introduced into the rotor blades. For example, fluctuations in the flow of a working fluid being compressed or in hot combustion gases or steam may cause the rotor blades to vibrate. One of the fundamental design considerations for turbomachinery designers is to control high cycle fatigue of the rotor blades by avoiding or minimizing resonance at their natural frequencies, as well as dynamic stresses caused by forced response and / or aeroelastic instabilities.

[0004] To improve the high cycle fatigue life of rotor blades, vibration dampers are typically provided under and / or between the platforms to frictionally dissipate vibration energy and reduce the amplitude of corresponding vibrations during operation. The amount of vibration energy removed by a vibration damper is a function of the dynamic weight and reaction load of the vibration damper.

[0005] While known dampers are adequate for typical operations, there is a desire to improve the overall damper effectiveness. For example, one problem with many known dampers is that they provide damping only at isolated contact locations with the associated rotor blade, such as at the ends and sometimes at central locations. Another problem with many known dampers is their inability to conform to complex platform shapes, such as shapes that include curved portions.

[0006]

[0005] Accordingly, improved damper designs are desired in the art. In particular, a damper design that provides improved damping throughout the length of the damper, for example, by increasing rotor blade contact, would be advantageous. Additionally, a damper design that conforms to complex platform geometries would be advantageous. Summary of the Invention

[0007] Aspects and advantages of the damper stack, rotor blades, and turbomachine according to the present disclosure will be set forth in part in the description which follows, or will be obvious from the description, or may be learned by practice of the present teachings.

[0008] According to one embodiment, a rotor blade for a turbomachine is provided. The rotor blade includes a body including a shank and an airfoil extending radially outward from the shank. The rotor blade further includes a platform surrounding the body, the platform including a slash face. The rotor blade further includes a damper stack disposed on the slash face and extending generally along the axial direction. The damper stack includes a plurality of damper pins, each of the plurality of damper pins contacting an adjacent damper pin.

[0009] According to another embodiment, a turbomachine is provided. The turbomachine includes a compressor section, a combustor section, and a turbine section. The turbomachine further includes a plurality of rotor blades disposed in at least one of the compressor section or the turbine section. Each of the plurality of rotor blades includes a body including a shank and an airfoil extending radially outward from the shank. Each of the plurality of rotor blades further includes a platform surrounding the body, the platform including a slash face. Each of the plurality of rotor blades further includes a damper stack disposed on the slash face and extending generally along the axial direction. The damper stack includes a plurality of damper pins, each of the plurality of damper pins contacting adjacent damper pins.

[0010] These and other features, aspects, and advantages of the present damper stack, rotor blade, and turbomachine will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the present technology and, together with the description, serve to explain the principles of the technology.

[0011] A full and enabling disclosure of the present damper stack, rotor blade, and turbomachine, including the best mode of making and using the present systems and methods, directed to one of ordinary skill in the art, is set forth in this specification, which makes reference to the accompanying drawings. [Brief description of the drawings]

[0012] [Figure 1] 1 is a schematic diagram of a turbomachine according to an embodiment of the present disclosure; [Diagram 2] 1 is a perspective view of a rotor blade according to an embodiment of the present disclosure; FIG. [Diagram 3] FIG. 4 is a perspective view of a rotor blade according to another embodiment of the present disclosure. [Figure 4] FIG. 2 is a side view illustrating adjacent rotor blades according to an embodiment of the present disclosure. [Diagram 5]FIG. 2 is a perspective view of a damper stack according to an embodiment of the present disclosure. [Figure 6] FIG. 2 is a cross-sectional view of a damper stack according to an embodiment of the present disclosure. [Figure 7] FIG. 13 is a cross-sectional view of a damper stack according to another embodiment of the present disclosure. [Figure 8] FIG. 13 is a cross-sectional view of a damper stack according to yet another embodiment of the present disclosure. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0013] Reference will now be made in detail to embodiments of the present damper stack, rotor blade, and turbomachine, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation of the present technology, not as a limitation of the present technology. Indeed, it will be apparent to those skilled in the art that modifications and variations can be made in the present technology without departing from the scope or spirit of the claimed technology. For example, features illustrated or described as part of one embodiment can be used in another embodiment to yield still further embodiments. Thus, the present disclosure is intended to cover such modifications and variations within the scope of the appended claims and their equivalents.

[0014] The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and description are used to refer to like or similar parts of the 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 denote the location or importance of the individual components.

[0015] Referring now to the drawings, Figure 1 shows a schematic diagram of one embodiment of a turbomachine, which in the illustrated embodiment is a gas turbine 10. Although an industrial or land-based gas turbine is shown and described herein, the present disclosure is not limited to land-based and / or industrial gas turbines unless otherwise stated in the claims. For example, the damping techniques described herein may be used with any type of turbomachine, including, but not limited to, a steam turbine, an aircraft gas turbine, or a marine gas turbine.

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

[0017] Compressor section 14 may generally include a number of rotor disks 24 (one of which is shown) and a number of rotor blades 26 extending radially outward from and connected to each rotor disk 24. Each rotor disk 24 may in turn be coupled to or form part of a portion of a shaft 22 that extends through compressor section 14.

[0018] The turbine section 18 may generally include a number of rotor disks 28 (one of which is shown) and a number of rotor blades 30 extending radially outward from and connected to each rotor disk 28. Each rotor disk 28 may in turn be coupled to or form a part of a portion of a shaft 22 that extends through the turbine section 18. The turbine section 18 further includes an outer casing 31 circumferentially surrounding the portion 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, enters the compressor section 14 through the inlet section 12 where the air is progressively compressed, thereby providing pressurized air to the combustors in the combustion section 16. The compressed air is mixed with fuel and combusted in each combustor to generate combustion gases 34. The combustion gases 34 enter the turbine section 18 from the combustor section 16 through the hot gas path 32, where energy (kinetic and / or thermal energy) is transferred from the combustion gases 34 to the rotor blades 30 to rotate the shaft 22. The mechanical rotational energy can then be used to power the compressor section 14 and / or generate electricity. The combustion gases 34 exhausted from the turbine section 18 may then be exhausted from the gas turbine 10 via the exhaust section 20.

[0020] 2 and 3 show an embodiment of a rotor blade according to an embodiment of the present disclosure. In the embodiment shown, the rotor blade is a turbine blade or bucket 30, but in alternative embodiments, the rotor blade may be a compressor blade or bucket 26.

[0021] The rotor blade 30 may include a body including an airfoil 36 and a shank 38. The airfoil 36 may be positioned extending radially outward from the shank 38. The shank 38 may include a root or dovetail 40 that may be attached to the rotor disk 28 to facilitate rotation of the rotor blade 30.

[0022] The airfoil 36 may have a generally aerodynamic profile. For example, the airfoil 36 may have an exterior surface defining a pressure side and a suction side, each extending between a leading edge and a trailing edge. The exterior surface of the shank 38 may include a pressure side, a suction side, a leading surface, and a trailing surface.

[0023] The platform 42 may generally surround the body. A typical platform may be positioned at an intersection or transition between the airfoil 36 and the shank 38 as shown and may extend generally axially and tangentially outward. In the turbine section 18, the platform 42 generally serves as a radially inward flow boundary for the combustion gases 34 flowing through the hot gas path 32. The platform 42 may also include axially spaced apart leading and trailing edge faces 52, 54. The leading edge face 52 is positioned in the flow of the combustion gases 34 and the trailing edge face 54 is positioned downstream from the leading edge face 52. Additionally, the platform 42 may include circumferentially spaced apart pressure side and suction side slashfaces 56, 58.

[0024] In some embodiments, the pressure side slashface 56 and / or the suction side slashface 58 may be generally planar (although may be conventionally planar or angled), as shown in Figure 2. In other embodiments, the pressure side slashface 56 and / or the suction side slashface 58, or at least a portion thereof, may be curviplanar, as shown in Figure 3. For example, the slashfaces 56 and / or 58 may be curved axially, radially, and / or tangentially.

[0025] As mentioned above, a plurality of rotor blades 30 may be provided on each of one or more rotor disks 28 and may extend radially outward therefrom. The rotor blades 30 provided on the rotor disks 28 may be assembled in a circumferential array such that the pressure side slashface 56 of each rotor blade 30 opposes the suction side slashface 58 of each adjacent rotor blade 30 when the rotor blades 30 are so assembled. FIG. 4 illustrates a pair of circumferentially adjacent adjacent rotor blades 30', 30''. As shown, the pressure side slashface 56 of the rotor blade 30'' opposes the suction side slashface 58 of the adjacent rotor blade 30' when the rotor blades 30', 30'' are so positioned.

[0026] 2-8, in accordance with the present disclosure, one or more damper stacks 70 may be provided in a groove 60 in a rotor blade 30. Specifically, the groove 60 may be defined in the slash faces 56, 58 of the platform 42 of the rotor blade 30. The groove 60 may extend generally along an axial direction. The damper stack 70 may be disposed within the groove 60 as shown in FIGS. 2 and 3. Notably, in an exemplary embodiment, a widthwise portion of the damper stack 70 protrudes from the groove 60 such that the damper stack 70 also contacts grooves 60 defined in adjacent slash faces 56, 58 of adjacent rotor blades 30 when assembled as discussed herein.

[0027] Each damper stack 70 may be disposed within and in contact with a slashface 56, 58 (e.g., pressure side slashface 56 or suction side slashface 58) of a rotor blade 30 and may extend generally axially and thus generally along the length of the respective slashface 56, 58, as shown. Additionally, as shown in FIG. 4 , a damper stack 70 according to the present disclosure may be disposed between and in contact with adjacent opposing pressure side slashfaces 56 or suction side slashfaces 58 of adjacent circumferentially-adjacent rotor blades 30.

[0028] A damper stack 70 according to the present disclosure advantageously functions as a vibration damper. During operation, the damper stack 70 frictionally dissipates vibration energy and reduces the amplitude of the corresponding vibrations. The amount of vibration energy removed by the damper stack 70 is a function of several factors, including, but not limited to, the dynamic weight of the damper stack 70, the geometry of the stack 70, and the reaction loads between adjacent rotor blades 30', 30''.

[0029] Each damper stack 70 may include multiple damper pins 72 arranged in an end-to-end arrangement corresponding to the linear or curvilinear shape of the groove 60. Each damper pin 72 may contact adjacent damper pins 72 in the damper stack 70. The use of damper stacks 70 in accordance with the present disclosure advantageously provides improved damping of rotor blades 30 in accordance with the present disclosure. For example, the use of multiple individual damper pins 72 in the arrangements shown and described herein advantageously provides improved damping throughout the length of the damper stack 70 due to the individual movement possible by each damper pin 72 of the damper stack 70 in the groove 60 and by the simultaneous contact of the damper stack 70 with the groove 60 of an adjacent rotor blade 30', 30''. Additionally, such damper stacks 70 and their associated damper pins 72 may conform to complex planar, curved, and / or partially curved platform shapes.

[0030] 5-8, various embodiments of a damper stack 70 according to the present disclosure are shown. As discussed, the damper stack 70 includes a plurality of damper pins 72. Each damper pin 72 has a length 73 defined between a first end 74 and a second end 76 of the damper pin 72. The damper pins 72 may be arranged in a linear array along their length such that adjacent ends 74, 76 of adjacent damper pins 72 contact one another.

[0031] For example, the multiple damper pins 72 may include a first damper pin 72′ and a second damper pin 72″, each of which extends between a first end 74 and a second end 76. The first end 74 of the first damper pin 72′ may contact the second end 76 of the second damper pin 72″. In some embodiments, the second end 76 of the first damper pin 72′ may contact another adjacent damper pin 72 and / or the first end 74 of the second damper pin 72″ may contact yet another adjacent damper pin 72.

[0032] In addition to contact with the rotor blades 30′, 30″, contact between adjacent damper pins 72 may provide a significant damping mechanism for the damping of the rotor blades 30′, 30″. Thus, the ends 74, 76 of adjacent damper pins 72 may have an appropriate shape to provide such primary damping. In some embodiments, the contacting ends 74, 76 of adjacent damper pins 72 may have complementary spherical shapes. For example, as shown in FIGS. 6 and 7 , the first end 74 of the first damper pin 72′ may have an outward (convex) spherical shape and the second end 76 of the second damper pin 72″ may have an inward (concave) spherical shape, or vice versa. In other embodiments, the contacting ends 74, 76 of adjacent damper pins 72 may have mirrored shapes. For example, as shown in Figure 8, the first end 74 of the first damper pin 72' and the second end 76 of the second damper pin 72'' may be flat surfaces that abut one another. Other suitable end 74, 76 shapes (such as conical, domed, or other complementary shapes) may be utilized provided the shapes provide adequate primary damping.

[0033] The damper pins 72 may have a generally elliptical or circular cross-sectional profile in some illustrated embodiments. Alternatively, cross-sectional profiles of other suitable shapes may be utilized. The cross-sectional profile may be constant or may vary along the length 73 of the damper pins 72. Furthermore, the damper pins 72 may have any suitable cross-sectional size. Still further, the damper pins 72 may be formed from any suitable material. The shape, size, and / or material may be the same for multiple damper pins 72 in the damper stack 70 or may vary for one or more of the damper pins 72 in the damper stack 70.

[0034] As discussed, each of the multiple damper pins 72 may have a length 73. In some embodiments, the lengths 73 of the damper pins 72 in the damper stack 70 may be the same. For example, as shown in FIGS. 6 and 8, the lengths 73 of the first and second damper pins 72', 72'' may be the same. In other embodiments, as shown in FIG. 7, the length 73 of one or more damper pins 72 in the damper stack 70 may differ from the other damper pins 72 in the stack. For example, as shown in FIG. 7, the length 73 of the first damper pin 72' may be different from the length 73 of the second damper pin 72''.

[0035] In the exemplary illustrated embodiment, each of the plurality of damper pins 72 has a hollow cross-sectional profile such that an internal passageway 78 is defined through the damper pins 72 and the damper stack 70. In other embodiments, damper pins 72 and damper stacks 70 according to the present disclosure may be solid such that no internal passageway is defined therethrough.

[0036] Additionally, in some embodiments shown, a wire 80 may extend through one or more damper pins 72 of the damper stack 70. For example, the wire 80 may extend through the interior passage 78 or through a separately defined interior passage, thus leaving the passage 78 empty. The wire 80 may generally join the damper pins 72 together. In other embodiments, other suitable components may be utilized to join the damper pins 72 together, or the damper pins 72 may not be joined together.

[0037] The damper stack 70 may include end pins 82 that are the outermost damper pins 72 at each end of the damper stack 70. The end pins 82 are spaced apart from one another along an axial direction. As shown, the damper stack 70 includes a first end pin 82' and a second end pin 82'' that are spaced apart along an axial direction. In some embodiments shown, a cutout portion 84 may be defined in each such end pin 82', 82'', such as at the outermost section of the pin 82', 82'' that includes the outermost end 74 or 76, i.e., the end that is not adjacent to another pin 72. This cutout portion 84 defines a shoulder 86 of the end pin 82', 82''. The shoulder 86 may include a support surface 88, which in the exemplary embodiment may be a flat plane.

[0038] In some embodiments, the groove 60 includes one or more shoulder slot portions 62, such as at each end of the groove 60. Such portions 62 define a support surface 64, which in exemplary embodiments may be a flat plane. In these embodiments, the shoulders 86 of the end pins 82', 82'' may be disposed within such shoulder slot portions 62 such that the support surface 88 may contact the support surface 64. Thus, the damper stack 70 may be supported within the groove 60 and undesired rotation during use and operation may be reduced or prevented.

[0039] This specification uses examples to disclose the invention, including the best mode, and to enable any person skilled in the art to practice the invention, including making and using any device or system, and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements that do not differ insubstantial from the literal language of the claims. [Explanation of symbols]

[0040] 10. Gas Turbine 12 Entrance Section 14 Compressor Section 16 Combustor Section 18 Turbine Section 20 Exhaust Section 22 Shaft 24 rotor disk 26 Rotor Blades 28 Rotor disc 30 Rotor Blades 30' rotor blade 30'' rotor blade 31 Outer casing 32 Hot Gas Path 34 Combustion Gas 36 Airfoil 38 Shank 40 Base / Dovetail 42 Platform 52 Leading edge surface 54 Trailing edge surface 56 Pressure side slash face 58 Negative pressure side slash face 60 grooves 62 Shoulder slot area 64 Support surface 70 Damper stack 72 Damper pin 72' First damper pin 72'' 2nd damper pin 73 Length 74 First end 76 Second End 78 Internal passage 80 Wire 82 Endpin 82' 1st Endpin 82'' 2nd endpin 84 Cutout part 86 Shoulder 88 Support surface

Claims

1. A turbomachine, comprising: A compressor section (14); a combustor section (16); A turbine section (18); a plurality of rotor blades (26, 30) provided in at least one of the compressor section (14) or the turbine section (18); each of the plurality of rotor blades (26, 30) a body including a shank (38) and an airfoil (36) extending radially outward from the shank (38); a platform (42) surrounding said body, said platform (42) having slash faces (56, 58); a damper stack (70) disposed on the slash faces (56, 58) and extending generally along an axial direction, the damper stack (70) including a plurality of damper pins (72), each of the plurality of damper pins (72) contacting an adjacent damper pin (72); Equipped with each of the plurality of damper pins (72) extends between a first end (74) and a second end (76), a first end (74) of a first damper pin (72') contacts a second end (76) of a second damper pin (72''), the first end (74) of the first damper pin (72') having a convex spherical shape and the second end (76) of the second damper pin (72'') having a concave spherical shape.

2. 2. The turbomachine of claim 1, wherein a length (73) of each of the plurality of damper pins (72) is defined between the first end (74) and the second end (76) of the damper pin (72), and the length (73) of the first damper pin (72') is different from the length (73) of the second damper pin (72'').

3. The turbomachine of claim 1 or claim 2, wherein the damper stack (70) further comprises a wire (80) extending through each of the plurality of damper pins (72).

4. The turbomachine of any of the preceding claims, wherein a groove (60) is defined in the slashfaces (56, 58), and the damper stack (70) is partially disposed within the groove (60).

5. 5. The turbomachine of claim 1, wherein the plurality of damper pins (72) comprises a first end pin (82') and a second end pin (82''), the first end pin (82') spaced apart from the second end pin (82'') along the axial direction, the first end pin (82') and the second end pin (82') each comprising a shoulder (86) defined by a cut-out portion (84).

6. The turbomachine of claim 5 , wherein each of the shoulders (86) includes a flat support surface (64).

7. 7. The turbomachine of claim 5 or claim 6, wherein a groove (60) is defined in the slash face, the damper stack (70) is disposed partially within the groove (60), and each of the shoulders (86) is disposed within a shoulder slot portion (62) of the groove (60).

8. The turbomachine of any preceding claim, wherein the plurality of rotor blades (30) are provided in the turbine section (18).

9. The turbomachine of any one of claims 1 to 8, wherein the turbomachine is a gas turbine (10).

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