Improved rotor blade vibration damping structure

Abstract

To provide improved rotor blade platform designs that provide decreased stiffness while still providing required structural integrity for a blade, and that allow the use of both a vibration damper and a platform seal.SOLUTION: Rotor blades 30 and turbomachines are provided. The rotor blades 30 include a main body 35 having a shank 38, an airfoil 36 extending radially outwardly from the shank 38, and a platform 42. The main body 35 further includes a pressure-side slash face 56 and a suction-side slash face 58. Each of the pressure-side slash face 56 and the suction-side slash face 58 includes a damper land 120 and defines a slot 70. The damper land 120 is disposed radially inward from the slot 70.SELECTED DRAWING: Figure 3

Description

[Technical Field]

[0001] FIELD OF THE DISCLOSURE This disclosure relates generally to rotor blades for turbomachines, and more particularly to improved rotor blade vibration damping structures. [Background technology]

[0002] Turbomachines are utilized in various industries and applications for the purpose of energy transfer. For example, a gas turbine engine typically 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 produce 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 then exit the gas turbine through 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 controlling high-cycle fatigue of the rotor blades by avoiding or minimizing resonance at their natural frequencies and dynamic stresses caused by forced response and / or aeroelastic instabilities.

[0004] For example, to improve the high cycle fatigue life of rotor blades, vibration dampers are typically provided below and / or between the platforms to frictionally dissipate vibration energy and reduce the amplitude of corresponding vibrations during operation.

[0005] The use of vibration dampers on known rotor blade platforms presents challenges. The design of the rotor blade platform directly impacts the effectiveness of the vibration damper during operation. For example, one known challenge is that the stiffness of known blade platforms, required to maintain structural integrity, results in low vibration damping effectiveness. Another challenge with many known blade platforms is the limited space available on the platform itself for installing vibration dampers. For example, the use of vibration dampers on the blade platform may limit or prevent the use of leak-tight seals due to a lack of space.

[0006] Therefore, there is a need in the art for improved rotor blade platform designs. In particular, there is a need for platforms that reduce stiffness while still providing the necessary structural integrity for the blades. Furthermore, there is a need for rotor blade platform designs that allow for the use of both vibration dampers and platform seals. Summary of the Invention

[0007] Aspects and advantages of the rotor blade and turbomachine according to the present disclosure will be set forth in part in the description that follows, or will be obvious from the description, or may be learned through practice of the present teachings.

[0008] According to one embodiment, a rotor blade for a turbomachine is provided. The rotor blade includes a body having a shank, an airfoil extending radially outward from the shank, and a platform. The body further includes a pressure side slashface and a suction side slashface. Each of the pressure side slashface and the suction side slashface includes a damper land and defines a slot. On both the pressure side slashface and the suction side slashface, the damper land is located radially inward from the slot.

[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 provided in at least one of the compressor section or the turbine section. Each of the plurality of rotor blades includes a body having a shank, an airfoil extending radially outward from the shank, and a platform. The body includes a pressure side slashface and a suction side slashface. Each of the pressure side slashface and the suction side slashface includes a damper land and defines a slot. On both the pressure side slashface and the suction side slashface, the damper land is located radially inward from the slot.

[0010] These and other features, aspects, and advantages of the present rotor blades and turbomachines 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, directed to one skilled in the art, of making and using the present systems and methods, is set forth in this specification, which makes reference to the accompanying drawings. [Brief explanation of the drawings]

[0012] [Figure 1] 1 is a schematic diagram of a turbomachine according to an embodiment of the present disclosure; [Figure 2] FIG. 2 is a pressure side perspective view of a rotor blade according to an embodiment of the present disclosure. [Figure 3] FIG. 2 is a suction side perspective view of a rotor blade and damper pin according to an 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. [Figure 5] FIG. 2 is an enlarged perspective view of the pressure side of a rotor blade according to an embodiment of the present disclosure. [Figure 6] FIG. 4 is an enlarged perspective view of the pressure side of a rotor blade according to another embodiment of the present disclosure. [Figure 7] FIG. 10 is an enlarged perspective view of the pressure side of a rotor blade according to yet a further embodiment of the present disclosure. [Figure 8] FIG. 2 is an enlarged cross-sectional view showing damper lands of two adjacent rotor blades according to an embodiment of the present disclosure. DETAILED DESCRIPTION OF THE INVENTION

[0013] Reference will now be made in detail to the present rotor blade and turbomachine embodiments, 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 thereof. Indeed, it will be apparent to those skilled in the art that various 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 a further embodiment. Accordingly, the present disclosure is intended to cover such modifications and variations as come 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" can be used interchangeably to distinguish one component from another and are not intended to denote the location or importance of individual components.

[0015] As used herein, the terms "upstream" (or "forward") and "downstream" (or "aft") refer to relative directions with respect to fluid flow in a fluid pathway. For example, "upstream" refers to the direction from which fluid flows, and "downstream" refers to the direction from which fluid flows. The term "radially" refers to relative directions that are substantially perpendicular to the axial centerline of a particular component, the term "axially" refers to relative directions that are substantially parallel and / or coaxially aligned with the axial centerline of a particular component, and the term "circumferentially" refers to relative directions that extend around the axial centerline of a particular component.

[0016] 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 vertical" includes any direction, e.g., clockwise or counterclockwise, within 10 degrees of vertical.

[0017] Referring now to the drawings, Figure 1 shows a schematic diagram 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 invention 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.

[0018] As shown, the gas turbine 10 generally includes an inlet section 12, a compressor section 14 disposed downstream from the inlet section 12, a plurality 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.

[0019] 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 and connected to each rotor disk 24. Each rotor disk 24 may in turn be coupled to or form part of a shaft 22 that extends through compressor section 14.

[0020] 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 and connected to each rotor disk 28. Each rotor disk 28 may then be coupled to or form part of a portion of a shaft 22 that extends through turbine section 18. Turbine section 18 further includes an outer casing 31 circumferentially surrounding the portion of shaft 22 and rotor blades 30, thereby at least partially defining a hot gas path 32 through turbine section 18.

[0021] During operation, a working fluid, such as air, enters the compressor section 14 through the inlet section 12, where it is progressively compressed, thereby providing compressed air to one or more combustors in the combustor 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, causing the shaft 22 to rotate. The 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 may then be exhausted from the gas turbine 10 through the exhaust section 20.

[0022] 2-8 illustrate embodiments of rotor blades according to embodiments of the present disclosure. In the illustrated embodiment, the rotor blades are turbine blades or rotor blades 30, but in alternative embodiments, the rotor blades may be compressor blades or rotor blades 26.

[0023] The rotor blade 30 may include a body 35 including an airfoil 36 and a shank 38. The airfoil 36 may be disposed 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.

[0024] The airfoil 36 may have a generally aerodynamic profile. For example, the airfoil 36 may have an outer surface defining a pressure side and a suction side, each extending between a leading edge and a trailing edge. The outer surface of the shank 38 may include a pressure side, a suction side, a leading surface, and a trailing surface.

[0025] The body 35 may further include a platform 42. A typical platform may be located at the intersection or transition between the airfoil 36 and the shank 38, as shown, and may extend generally axially and tangentially outward relative to the shank. In the turbine section 18, the platform 42 generally serves as a radially inner flow boundary for the combustion gases 34 flowing through the hot gas path 32. The platform 42 may include a leading edge surface 52 spaced axially from a trailing edge surface 54. The leading edge surface 52 is disposed in the flow of the combustion gases 34, and the trailing edge surface 54 is disposed downstream from the leading edge surface 52. Additionally, the body 35 may include a pressure side slashface 56 spaced circumferentially from a suction side slashface 58.

[0026] In some embodiments, as shown in Figures 2 and 3, the pressure side slashface 56 and / or the suction side slashface 58 may be a generally planar surface (conventionally, sometimes planar or curved). In other embodiments, such as those shown in Figures 5-7, the pressure side slashface 56 and / or the suction side slashface 58, or at least portions thereof, may be curviplanar. For example, the slashfaces 56 and / or 58 may be curved axially, radially, and / or tangentially.

[0027] FIG. 4 illustrates a pair of circumferentially adjacent, adjacent rotor blades 30′, 30″. As shown, the pressure side slashface 56 of one rotor blade 30 opposes the suction side slashface 58 of the adjacent rotor blade 30 when the rotor blades 30 are so arranged. As described above, a plurality of rotor blades 30 may be provided on one or more respective rotor disks 28 and extend radially outward therefrom. The rotor blades 30 provided on the rotor disks 28 are 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. In some embodiments, the pressure side slashface 56 of each rotor blade 30 and the suction side slashface 58 of each adjacent rotor blade 30 may define a circumferential gap 60.

[0028] 3 , in accordance with the present disclosure, one or more damper pins 95 may be provided on the rotor blade 30. Each damper pin 95 may include a first end 200 axially spaced from a second end 202. The first end 200 and the second end 202 may include shoulders 204, 206, respectively. Each damper pin 95 may be positioned on and in contact with a slashface 56, 58 (e.g., the pressure side slashface 56 or the suction side slashface 58) of the rotor blade 30 and may extend generally axially and, therefore, generally along the length of the slashface 56, 58, as shown.

[0029] Additionally, as shown in FIG. 4 , a damper pin 95 according to the present disclosure may be positioned between and contact adjacent opposing pressure side slashfaces 56 or suction side slashfaces 58 of adjacent circumferentially adjacent rotor blades 30.

[0030] The damper pin 95 according to the present disclosure advantageously functions as a vibration damper. During operation, the damper pin 95 frictionally dissipates vibration energy, reducing the amplitude of the corresponding vibrations.

[0031] 2 and 3 illustrate the pressure side slashface 56 and the suction side slashface 58 of the body 35. As shown, the body 35 may include one or more slots 70 defined in the pressure side slashface 56 and / or the suction side slashface 58 of the body 35. In some embodiments, the slot 70 may be a continuous groove defined along each of the pressure side slashface 56 and the suction side slashface 58. The slot 70 may include a leading edge segment 72, a platform segment 74, and a trailing edge segment 76. The leading edge segment 72 may be defined along the leading surface 52, the platform segment 74 may be defined along the platform 42, and the trailing edge segment 76 may be defined along the trailing surface 56. As used herein, terms such as "defined along" and cognates may mean "substantially parallel to" or "generally aligned with."

[0032] In other embodiments, the leading edge segment 72 and the trailing edge segment 76 of the slot 70 may be oriented generally radially with respect to the axial centerline of the gas turbine 10. Similarly, the platform segment 74 of the slot 70 may be oriented generally axially with respect to the axial centerline of the gas turbine 10. In some embodiments, the leading edge segment 72 may be directly connected to and contiguous with the platform segment 74, and the platform segment 74 may be directly connected to and contiguous with the trailing edge segment 76. In some embodiments, the platform segment 74 may be defined within the platform 42 and be oriented axially with respect to the axial centerline of the gas turbine 10.

[0033] In an alternative embodiment (not shown), the slot 70 may be discontinuous. In such an embodiment, the leading edge segment 72, the platform segment 74, and the trailing edge segment 76 may be entirely separate slots or grooves defined circumferentially in the pressure side slashface 56 and / or the suction side slashface 58.

[0034] As shown in FIG. 3 , the body 35 may further include a suction side damper land 120. The suction side damper land 120 may include a first end 122 axially spaced from a second end 124. In many embodiments, a cutout 121 may be defined in the suction side damper land 120. The cutout 121 includes a shoulder slot portion 126 formed in the first end 122 and the second end 124 of the suction side damper land 120. The shoulder slot portion 126 defines a support surface 128, which may be a flat plane in exemplary embodiments. In these embodiments, the shoulders 204 and 206 of the damper pin 95 may be positioned in the shoulder slot portion 126 such that the support surface 128 may contact the shoulders 204 and 206. Thus, the damper pin 95 may be supported within the suction side damper land 120, reducing or preventing undesirable over-rotation during use and operation.

[0035] 3 , the slot 70 may be sized to securely contain a portion of the seal 84 therein, i.e., the slot 70 may be sized to prevent the seal 84 from slipping out of the slot 70 during operation of the gas turbine 10. The seal 84 may include and extend between a first end 86 and a second end 88. The seal 84 may be sized to at least partially sealingly fit within the slot 70.

[0036] When two or more blades 30 are positioned adjacent to one another on the rotor disk 24, as in the configurations shown in FIGS. 4 and 8 and described above, the slots 70 in the pressure side slashface 56 of each rotor blade 30 align with the slots 70 in the suction side slashface 58 of an adjacent rotor blade 30 to define a channel. The rotor blades 30 positioned adjacent to one another may include rotor blades 30 directly adjacent to one another on the rotor disk 24 and / or rotor blades 30 in direct contact with one another. A seal 84 (shown in FIG. 3) may be housed within the channel defined by each slot 70. The seal 84 may extend between and into the slots 70 of both adjacent rotor blades 30′, 30″. In some embodiments, the seal 84 prevents unwanted hot gases from the turbine section 18 from leaking into the body 35 of the blade 30. Alternatively, or additionally, in many embodiments, the seal 84 may prevent compressed cooling air from the compressor section 14 from leaking from the shank 38 into the turbine section 18 .

[0037] As shown in FIGS. 5-7 , the pressure side slashface 56 may further include a pressure side damper land 90 having a first end 92 and a second end 94. In some embodiments, the first end 92 of the pressure side damper land 90 may be axially spaced from the second end 94. In various embodiments, the first end 92 of the pressure side damper land 90 may partially define the leading edge segment 72 of the slot 70 and may extend to the second end 94, which partially defines the trailing edge segment 76 of the slot 70. In many embodiments, the pressure side damper land 90 may be positioned radially inward from the slot 70. Specifically, the pressure side damper land 90 may be positioned radially inward from the platform segment 74 of the slot 70. In some embodiments, the pressure side damper land 90 may be oriented parallel to the platform segment 74 of the slot 70. In other embodiments, the pressure side damper land 90 may be oriented axially relative to the axial centerline of the gas turbine 10.

[0038] The pressure side damper land 90 may function to provide a surface for positioning a damper pin 95 thereon and to provide vibration damping to the rotor blade 30. In many embodiments, the surface of the pressure side damper land 90 may be shaped to slightly match the shape of the damper pin 95 to enhance surface contact and vibration damping. For example, the pressure side damper land 90 further includes a curved portion 96 and a flat portion 98. The curved portion 96 may curve circumferentially inward from the platform 42 to the flat portion 98. The flat portion 98 of the pressure side damper land 90 may extend radially inward from the curved portion 96 to the shank cutout 39 defined in the body 35. The flat portion 98 may be generally parallel to the platform 42 in both the axial and radial directions of the gas turbine 10.

[0039] In some embodiments, the pressure side damper land 90 may be substantially cantilevered due to the slot 70 and its flat portion 98. For example, the flat portion 98 of the pressure side damper land 90 may extend radially inward from the curved portion 96 to a free end 99. The free end 99 may be substantially cantilevered within the shank cutout 39 to advantageously increase the overall compliance of the platform 42 of the rotor blade 30, thereby effectively increasing vibration damping. In various embodiments, the flat portion 98 of the pressure side damper land 90 may taper axially inward from the curved portion 96 to the free end 99. In various embodiments, the flat portion 98 of the pressure side damper land 90 may taper from a first undercut 100 and a second undercut 102 at each end.

[0040] In addition to providing a housing for the seal 84, the slot 70 may provide reduced material stiffness and increased compliance in the pressure side damper land 90, enabling increased vibration damping for the entire blade 30. Additionally, the slot 70 may include a slot depth 71. Varying, i.e., increasing or decreasing, the slot depth 71 advantageously increases or decreases the overall stiffness of the pressure side damper land 90, thereby increasing the overall damping effectiveness.

[0041] In some embodiments, such as those illustrated in FIGS. 5-7, the pressure side slashface 56 may further include a first undercut 100 and a second undercut 102. The first undercut 100 and the second undercut 102 advantageously function to modify, i.e., increase or decrease, the stiffness of the shank 38 to enhance overall vibration damping. In the embodiment illustrated in FIG. 5, the first undercut 100 and the second undercut 102 may be semicircular cuts defined circumferentially inward on the body 35 of the rotor blade 30. In some embodiments, the first undercut 100 and the second undercut 102 may be substantially curved or arc-shaped. In other embodiments, such as those illustrated in FIGS. 6 and 7, the first undercut 100 and the second undercut 102 may include multiple semicircular cuts or trapezoidal cuts.

[0042] In many embodiments, the first undercut 100 may be located directly radially inward from the first end 92 of the pressure side damper land 90, and the second undercut 102 may be axially spaced from the first undercut 100 and may be located directly radially inward from the second end 94 of the pressure side damper land 90. In some embodiments, the first undercut 100 and the second undercut 102 may extend generally radially inward from the first end 92 and the second end 94 of the pressure side damper land 90, respectively. In many embodiments, both the first undercut 100 and the second undercut 102 may extend radially inward beyond the free end 99 of the pressure side damper land 90.

[0043] The first undercut 100 and the second undercut 102 may each partially define the slot 70. More specifically, the first undercut 100 may partially define the leading edge segment 72 of the slot 70. Similarly, the second undercut 102 may partially define the trailing edge segment 76 of the slot 70. In various embodiments, the first undercut 100 may be axially disposed between the leading edge segment 72 of the slot 70 and the flat portion 98 of the pressure side damper land 90. The second undercut 102 may be axially disposed between the flat portion 98 of the pressure side damper land 90 and the trailing edge segment 76 of the slot 70.

[0044] The first undercut 100 and the second undercut 102 may each have a maximum undercut depth 106 defined in the circumferential direction. The maximum undercut depth 106 of the first undercut 100 may be the same as or different from the maximum undercut depth 106 of the second undercut 102. Varying the maximum undercut depth 106 of the first undercut and / or the second undercut 102 advantageously changes, i.e., increases or decreases, the stiffness of the pressure side damper land 90, enhancing damping effectiveness. In some embodiments, the maximum undercut depth 106 of each undercut 100, 102 may be up to approximately 1.5 inches. In other embodiments, the maximum undercut depth 106 may be up to approximately 1 inch. In some embodiments, the maximum undercut depth 106 may be up to approximately 0.75 inches. In various embodiments, the maximum undercut depth 106 may be up to approximately 0.5 inches. In other embodiments, the maximum undercut depth 106 may be up to approximately 0.25 inches.

[0045] FIG. 8 illustrates a cross-sectional view of a pair of circumferentially adjacent rotor blades. As shown, the pressure side damper land 90 of a first rotor blade 30′ aligns with the suction side damper land 120 of the adjacent second rotor blade 30″ when the rotor blades 30′, 30″ are so positioned. As shown in FIG. 8 , a damper pin 95 may be positioned along the suction side damper land 120. During operation, the damper pin 95 may move in the direction of arrow 130 and contact both the pressure side damper land 90 and the suction side damper land 120 to provide vibration damping to the adjacent rotor blades 30′, 30″.

[0046] Additionally, the slot depth 71' of the pressure side slashface 56 may be different from the slot depth 71'' of the suction side slashface 58. For example, the slot depth 71' of the pressure side slashface 56 may be greater or less than the slot depth 71'' of the suction side slashface 58. Generally, the sum of the slot depth 71' of the pressure side slashface 56, the width of the circumferential gap 60 (shown in FIG. 4), and the slot depth 71'' of the suction side slashface 58 may be approximately equal to or slightly greater than the width of the seal 84. In various embodiments, the seal 84 may be smaller than the slot 70 to allow for thermal expansion within the slot 70. Additionally, the seal 84 may be sized to allow for manufacturing variations therein.

[0047] For example, in many embodiments, the width of the seal 84 can be from about 5% to about 30% of the width of the channel, allowing for both manufacturing variations and thermal expansion within the slot 70. The embodiments shown in Figures 2-8 allow for the use of both the vibration damping damper pin 95 and the seal 84. In various embodiments, the blade 30 can include only the vibration damping damper pin 95, only the seal 84, or both the vibration damping damper pin 95 and the seal 84.

[0048] This written description 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 contain structural elements that do not differ from the literal language of the claims, or if they contain equivalent structural elements that do not differ insubstantial way from the literal language of the claims. [Explanation of symbols]

[0049] 10. Gas turbine 12 Entrance Section 14 Compressor Section 16 Combustor Section 18 Turbine Section 20 Exhaust Section 22 shaft 24 rotor disc 26 Rotor blades, moving blades 28 rotor disc 30 Rotor blades, moving blades 30' First rotor blade 30'' second rotor blade 31 outer casing 32 Hot gas path 34 Combustion Gas 35 Main Unit 36 Airfoil 38 Shank 39 Shank notch 40 Root, dovetail 42 Platform 52 Leading edge surface 54 Trailing edge surface 56 Pressure side slash face 58 Suction side slash face 60 gap 70 slots 71 slot depth 71' slot depth 71'' slot depth 72 leading edge segments 74 Platform Segments 76 Trailing Edge Segment 84 stickers 86 First end 88 Second end 90 Pressure side damper land 92 first end 94 Second end 95 Vibration damping damper pin 96 curved part 98 Flat part 99 Free end 100 First Undercut 102 Second Undercut 106 Maximum undercut depth 120 Negative pressure side damper land 121 Notch 122 first end 124 Second End 126 Shoulder slot part 128 Support surface 130 Arrow 200 first end 202 second end 204 Shoulder 206 Shoulder

Claims

1. A rotor blade (30) for a turbomachine, the rotor blade (30) comprising: A body (35) having a shank (38), an airfoil (36) extending radially outward from the shank (38), and a platform (42), the body having a pressure side slashface (56), a suction side slashface (58), a leading edge surface, and a trailing edge surface. Equipped with each of the pressure side slashface (56) and the suction side slashface (58) includes a damper land (90), (120) and defines a slot (70); The damper land (90) on the pressure side slashface (56) is disposed radially inward from at least a portion of the slot (70) on the pressure side slashface (56), the damper land (90) on the pressure side slashface (56) and the slot (70) on the pressure side slashface (56) at least partially define one or more pressure side undercuts, and the damper land (120) on the suction side slashface (58) defines one or more pressure side undercuts. a damper land (90) on the pressure side slashface (56) and a damper land (120) on the suction side slashface (58) disposed radially inward from at least a portion of the slot (70) on the pressure side slashface (56), the damper land (90) on the pressure side slashface (56) and the damper land (120) on the suction side slashface (58) including a first end and a second end, the first end being axially spaced from the second end, the first end being near the leading surface and the second end being near the trailing surface.

2. 2. The rotor blade of claim 1, wherein the pressure side slashface further comprises one or more undercuts disposed radially inward of the first and second ends of the damper land of the pressure side slashface, each of the one or more undercuts having a circumferentially defined maximum undercut depth.

3. The rotor blade (30) of claim 2, wherein each of the one or more undercuts (100), (102) is arcuate.

4. The rotor blade (30) of claim 2 or 3, wherein the maximum undercut depth (106) of each of the one or more undercuts (100), (102) is a maximum of about 1.5 inches.

5. The rotor blade (30) of any one of claims 2 to 4, wherein the one or more undercuts (100), (102) at least partially define the slot (70) in the pressure side slash face (56).

6. 6. The rotor blade of claim 1, wherein both the pressure side slashface slot and the suction side slashface slot comprise a leading edge segment, a platform segment, and a trailing edge segment, the leading edge segment defined along the leading surface, the platform segment defined along the platform, and the trailing edge segment defined along the trailing surface.

7. 7. The rotor blade of claim 6, wherein the damper land on the pressure side slashface is disposed radially inward of the platform segment of the slot defined in the pressure side slashface, and the damper land on the suction side slashface is disposed radially inward of the platform segment of the slot defined in the suction side slashface.

8. a compressor section (14); a combustor section (16); a turbine section (18); a plurality of rotor blades (30) provided in at least one of the compressor section (14) or the turbine section (18), each of the plurality of rotor blades (30) comprising: A body (35) having a shank (38), an airfoil (36) extending radially outward from the shank (38), and a platform (42), the body having a pressure side slashface (56), a suction side slashface (58), a leading edge surface, and a trailing edge surface. Equipped with each of the pressure side slashface (56) and the suction side slashface (58) includes a damper land (90), (120), respectively, and defines a slot (70); The damper land (90) on the pressure side slashface (56) is disposed radially inward from at least a portion of the slot (70) on the pressure side slashface (56), the damper land (90) on the pressure side slashface (56) and the slot (70) on the pressure side slashface (56) at least partially define one or more pressure side undercuts, and the damper land (120) on the suction side slashface (58) is disposed radially inward from at least a portion of the slot (70) on the pressure side slashface (56), the damper land (90) on the pressure side slashface (56) and the slot (70) on the pressure side slashface at least partially define one or more pressure side undercuts. a plurality of rotor blades (30) disposed radially inward from at least a portion of the slots (70) on the pressure side slashface (56) and a damper land (120) on the suction side slashface (58), the damper land (90) on the pressure side slashface (56) and the damper land (120) on the suction side slashface (58) including a first end and a second end, the first end being axially spaced from the second end, the first end being near the leading edge surface and the second end being near the trailing edge surface; A turbomachine comprising:

9. 9. The turbomachine of claim 8, wherein the pressure side slashface further comprises one or more undercuts disposed radially inward of the first and second ends of the damper land of the pressure side slashface, each of the one or more undercuts having a circumferentially defined maximum undercut depth.

10. The turbomachine of claim 9, wherein each of the one or more undercuts (100), (102) is arcuate.

11. The turbomachine of claim 9 or 10, wherein the maximum undercut depth (106) of each of the one or more undercuts (100), (102) is up to about 1.5 inches.

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