Improved rotor blade damping structure

Abstract

The invention is entitled "Improved Rotor Blade Damping Structure". The invention provides a rotor blade (30) and a turbomachine. The rotor blade (30) includes a main body (35) having a shank (38), a platform (42), and an airfoil (36) extending radially outward from the shank (38). The main body (35) further includes a pressure side impingement surface (56) and a suction side impingement surface (58). Each of the pressure side impingement surface (56) and the suction side impingement surface (58) includes a damper blade band (90), (120) and defines a slot (70). The damper blade band (90), (120) is disposed radially inward from the slot (70).

Description

TECHNICAL FIELD

[0001] The present disclosure relates generally to rotor blades for turbomachinery, and more particularly, to improved rotor blade damping structures. BACKGROUND

[0002] Turbomachinery is used in a variety of 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 progressively increases the pressure of a working fluid entering the gas turbine engine and supplies this compressed working fluid to the combustion section. The compressed working fluid and a fuel, such as natural gas, are mixed and combusted in a combustion chamber within the combustion section to generate high pressure and temperature combustion gases. The combustion gases flow from the combustion section into the turbine section where the combustion gases expand to do work. For example, the expansion of the combustion gases in the turbine section can cause a rotor shaft connected to, for example, an electrical generator to rotate to produce electricity. The combustion gases then exit the gas turbine via the exhaust section.

[0003] The compressor section and turbine section generally include a plurality of rotor blades typically arranged in a plurality of stages. During engine operation, vibrations can be introduced into the rotor blades. For example, flow fluctuations of the compressed working fluid or hot combustion gases or steam can cause the rotor blades to vibrate. One fundamental design consideration for turbomachinery designers is to avoid or minimize resonance with the natural frequencies of the rotor blades and dynamic stresses resulting from forced response and / or aeroelastic instability, thereby controlling high cycle fatigue of the rotor blades.

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

[0005] There are problems with the use of vibration dampers in known rotor blade platforms. The design of the rotor blade platform directly impacts the effectiveness of the vibration dampers during operation. For example, one known problem is that the stiffness of known blade platforms required to maintain structural integrity results in lower vibration damping effectiveness. Another problem with many known blade platforms is that there is limited space on the platform itself to mount vibration dampers. For example, the use of vibration dampers on the blade platform can limit or inhibit the use of leak-proof seals due to lack of space.

[0006] Accordingly, there is a need in the art for improved rotor blade platform designs. In particular, it would be desirable to provide a platform that has reduced stiffness while still providing the required structural integrity for the blade. Furthermore, it would be desirable to have a rotor blade platform design that allows for the use of both vibration dampers and platform seals. SUMMARY

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

[0008] According to one embodiment, a rotor blade for a turbomachine is provided. The rotor blade includes a main body having a shank, a platform, and an airfoil extending radially outward from the shank. The main body further includes a pressure side impingement surface and a suction side impingement surface. Each of the pressure side impingement surface and the suction side impingement surface includes a damper blade band and defines a slot. In both the pressure side impingement surface and the suction side impingement surface, the damper blade band is disposed 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 disposed in at least one of the compressor section or the turbine section. Each rotor blade of the plurality of rotor blades includes a main body having a shank, a platform, and an airfoil extending radially outward from the shank. The main body includes a pressure side impingement surface and a suction side impingement surface. Each of the pressure side impingement surface and the suction side impingement surface includes a damper blade band and defines a slot. In both the pressure side impingement surface and the suction side impingement surface, the damper blade band is disposed radially inward from the slot.

[0010] These and other features, aspects, and advantages of the rotor blade and turbomachine of the present technology 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 technology and serve to explain the principles of the technology. BRIEF DESCRIPTION OF DRAWINGS

[0011] The complete and enabling disclosure of the damper stack, rotor blade, and turbomachine of the present technology, including the best mode thereof, directed to one of ordinary skill in the art in this field of technology is set forth in this specification, which is to be construed in conjunction with the appended drawings, wherein:

[0012] Figure 1 A schematic view of a turbomachine is shown in accordance with an embodiment of the present disclosure;

[0013] Figure 2 A pressure side perspective view of a rotor blade is shown in accordance with an embodiment of the present disclosure;

[0014] Figure 3 A suction side perspective view of a rotor blade and damper pin is shown in accordance with an embodiment of the present disclosure;

[0015] Figure 4 A side view showing adjacent rotor blades is shown in accordance with an embodiment of the present disclosure;

[0016] Figure 5 a magnified pressure side perspective view of a rotor blade is shown in accordance with other embodiments of the present disclosure;

[0017] Figure 6 a magnified pressure side perspective view of a rotor blade is shown in accordance with other embodiments of the present disclosure;

[0018] Figure 7 a magnified pressure side perspective view of a rotor blade is shown in accordance with yet other embodiments of the present disclosure; and

[0019] Figure 8 is a magnified cross-sectional view showing a damper blade band of two adjacent rotor blades in accordance with embodiments of the present disclosure. DETAILED DESCRIPTION

[0020] Reference will now be made in detail to embodiments of the rotor blade and turbomachine of the present application, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation of the present technology and is not meant as a limitation of the present technology. In fact, 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 technology as claimed. For instance, features illustrated or described as part of one embodiment, can be used with another embodiment to yield still a further embodiment. Thus, it is intended that this disclosure cover such modifications and variations as come within the scope of the appended claims and their equivalents.

[0021] The DETAILED DESCRIPTION uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the present application. As used herein, the terms "first", "second", and "third" can be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components.

[0022] As used herein, the terms "upstream" (or "up") and "downstream" (or "down") refer to the relative direction with respect to the flow of fluid in a fluid pathway. For example, "upstream" refers to the direction from which fluid flows, and "downstream" refers to the direction to which fluid flows. The term "radially" refers to a relative direction substantially perpendicular to an axial centerline of a particular component; the term "axially" refers to a relative direction substantially parallel and / or coaxially aligned with an axial centerline of a particular component; and the term "circumferentially" refers to a relative direction extending around an axial centerline of a particular component.

[0023] Approximating language, such as "generally" or "about," encompasses values and furnishings that are close to but not exactly a stated value. When referring to an angle or direction, such terms encompass angles that are within ten degrees of the stated angle or direction. For example, "generally vertical" encompasses directions that are within ten degrees of vertical in any direction (e.g., clockwise or counterclockwise).

[0024] Referring now to the drawings, Figure 1 A schematic view of a turbine, which in the illustrated embodiment is a gas turbine 10, is shown. 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 specified in the claims. For example, the present application as described herein can be used in any type of turbine, including but not limited to a steam turbine, an aircraft gas turbine, or a marine gas turbine.

[0025] As shown, the gas turbine 10 generally includes an inlet section 12, a compressor section 14 disposed downstream of the inlet section 12, a plurality of combustors (not shown) disposed within a combustor section 16 downstream of the compressor section 14, a turbine section 18 disposed downstream of the combustor section 16, and an exhaust section 20 disposed downstream of the turbine section 18. Additionally, the gas turbine 10 can include one or more shafts 22 coupled between the compressor section 14 and the turbine section 18.

[0026] The compressor section 14 can 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 can in turn be coupled to or can form a portion of the shaft 22 that extends through the compressor section 14.

[0027] The turbine section 18 can 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 interconnected to each rotor disk 28. Each rotor disk 28 can in turn be coupled to or can form a portion of the shaft 22 that extends through the turbine section 18. The turbine section 18 also includes an outer casing 31 that circumferentially surrounds 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.

[0028] During operation, a working fluid, such as air, flows through the inlet section 12 and into the compressor section 14, where the air is progressively compressed, providing pressurized air to one or more combustors of the combustion section 16. The pressurized air is mixed with fuel and combusted within each combustor to produce combustion gases 34. The combustion gases 34 flow from the combustor section 16 through the hot gas path 32, into the turbine section 18, where energy (kinetic and / or thermal) 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 exiting the turbine section 18 can then be exhausted from the gas turbine 10 via the exhaust section 20.

[0029] Figures 2 to 8 Embodiments of a rotor blade according to embodiments of the present disclosure are shown. In the illustrated embodiment, the rotor blade is a turbine blade or vane 30, although in alternative embodiments the rotor blade can be a compressor blade or vane 26.

[0030] The rotor blade 30 can include a main body 35 including an airfoil 36 and a shank 38. The airfoil 36 can extend and be positioned radially outward from the shank 38. The shank 38 can include a root or dovetail 40 that can be attached to the rotor disk 28 to facilitate rotation of the rotor blade 30.

[0031] The airfoil 36 can have a generally aerodynamic profile. For example, the airfoil 36 can 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 can include a pressure side face, a suction side face, a leading edge face, and a trailing edge face.

[0032] The main body 35 can also include a platform 42. A typical platform can be positioned at an intersection or transition between the airfoil 36 and the shank 38, and can extend outward relative to the shank in generally axial and tangential directions, as shown. 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 can include a leading edge face 52 axially spaced apart from a trailing edge face 54. The leading edge face 52 is positioned in the combustion gas flow 34, and the trailing edge face 54 is positioned downstream of the leading edge face 52. Further, the main body 35 can include a pressure side impingement face 56 circumferentially spaced apart from a suction side impingement face 58.

[0033] In some embodiments, as shown in FIGS. 1-3, the pressure side impingement face 56 and / or the suction side impingement face 58 can be a generally planar face (which can generally be planar or inclined). In other embodiments, such as shown in FIGS. 4-6, the pressure side impingement face 56 and / or the suction side impingement face 58 can be a generally curved face. Figure 2 and Figure 3 In some embodiments, as shown in FIGS. 1-3, the pressure side impingement face 56 and / or the suction side impingement face 58 can be a generally planar face (which can generally be planar or inclined). In other embodiments, such as shown in FIGS. 4-6, the pressure side impingement face 56 and / or the suction side impingement face 58 can be a generally curved face. Figures 5 to 7In the illustrated embodiment, the pressure side impact surface 56 and / or the suction side impact surface 58, or at least portions thereof, can be curvate planar. For example, the impact surfaces 56 and / or 58 can be curved with respect to the axial direction, the radial direction, and / or the tangential direction.

[0034] Figure 4 A pair of circumferentially adjacent adjacent rotor blades 30’, 30” are shown. As shown, when the rotor blades 30 are so positioned, the pressure side impact surface 56 of the rotor blade 30 faces the suction side impact surface 58 of the adjacent rotor blade 30. As noted above, a plurality of rotor blades 30 can be provided on each of one or more rotor disks 28, and can extend radially outwardly from each rotor disk. The rotor blades 30 provided on the rotor disk 28 can be assembled in a circumferential array such that, when the rotor blades 30 are so assembled, the pressure side impact surface 56 of each rotor blade 30 faces the suction side impact surface 58 of each adjacent rotor blade 30. In some embodiments, the pressure side impact surface 56 of each rotor blade 30 and the suction side impact surface 58 of each adjacent rotor blade 30 can define a gap 60 in the circumferential direction.

[0035] Referring again to Figure 3 , in accordance with the present disclosure, one or more damper pins 95 can be provided in the rotor blade 30. Each damper pin 95 can include a first end 200 axially separated from a second end 202. The first end 200 and the second end 202 can include shoulders 204, 206, respectively. Each damper pin 95 can be provided at and in contact with an impact surface 56, 58 (e.g., the pressure side impact surface 56 or the suction side impact surface 58) of the rotor blade 30, and can extend generally in the axial direction and thus generally along the length of the impact surface 56, 58, as shown.

[0036] Further, as Figure 4 shown, the damper pin 95 in accordance with the present disclosure can be provided between and in contact with adjacent, facing pressure side impact surfaces 56 or suction side impact surfaces 58 of circumferentially adjacent adjacent rotor blades 30.

[0037] The damper pin 95 in accordance with the present disclosure advantageously functions as a vibration damper. In operation, the damper pin 95 frictionally dissipates vibrational energy and reduces the corresponding amplitude of the vibration.

[0038] Figure 2 and Figure 3The pressure side impact surface 56 and the suction side impact surface 58 of the body 35 are shown. As shown, the body 35 may include one or more slots 70 defined in the pressure side impact surface 56 and / or the suction side impact surface 58 of the body 35. In some embodiments, the slot 70 may be one continuous groove defined along each of the pressure side impact surface 56 and / or the suction side impact surface 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 edge surface 52, the platform segment 92 may be defined along the platform 42, and the trailing edge segment 76 may be defined along the trailing edge surface 56. As used herein, terms such as "defined along..." and its cognates may mean "substantially parallel to..." or "generally in line with..."

[0039] In other embodiments, the leading edge segment 72 and the trailing edge segment 76 of the slot 70 can be oriented generally radially relative to the axial centerline of the gas turbine 10. Similarly, the platform segment 74 of the slot 70 can be oriented generally axially relative to the axial centerline of the gas turbine 10. In some embodiments, the leading edge segment 72 can be directly connected to and continuous with the platform segment 74, and the platform segment 74 can be directly connected to and continuous with the trailing edge segment 76. In some embodiments, the platform segment 74 can be defined within the platform 42 and oriented axially relative to the axial centerline of the gas turbine 10.

[0040] 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 completely separate slots or grooves circumferentially defined within the pressure side impact surface 56 and / or the suction side impact surface 58.

[0041] like Figure 3 As shown, the body 35 may also include a suction side damper land 120. The suction side damper land 120 may include a first end 122 axially separated from a second end 124. In many embodiments, a notch 121 may be defined in the suction side damper land 120. The notch 121 includes a shoulder slot portion 126 defined at 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, planar surface in exemplary embodiments. In these embodiments, the shoulders 204 and 206 of the damper pin 95 may be disposed 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 in the suction side damper land 120, and undesirable over-rotation during use and operation may be reduced or prevented.

[0042] In addition, if Figure 3As shown, the slot 70 can be sized to securely receive a portion of the seal 84 therein, i.e., the slot 70 can be sized to prevent the seal 84 from sliding out of the slot 70 during operation of the gas turbine 10. The seal 84 can include a first end 86 and a second end 88 and can extend therebetween. The seal 84 can be sized to at least partially sealingly fit into the slot 70.

[0043] When two or more blades 30 are arranged adjacent to each other on the rotor disk 24, such as Figure 4 and Figure 8 In the configuration shown and discussed above, the slots 70 of the pressure side impact surface 56 of each rotor blade 30 are aligned with the slots 70 of the suction side impact surface 58 of the adjacent rotor blade 30 to define a channel. The rotor blades 30 arranged adjacent to each other may include rotor blades 30 that are directly adjacent to each other on the rotor disk 24 and / or rotor blades 30 that are in direct contact with each other. The seal 84 (shown in FIG. Figure 3 70 . A seal 84 can extend between and into two slots 70 of 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 can prevent compressed cooling air from the compressor section 14 from leaking out of the shank 38 and into the turbine section 18.

[0044] like Figures 5 to 7 As shown, the pressure side impact surface 56 may also 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 separated from the second end 94. In various embodiments, the first end 92 of the pressure side damper land may partially define the leading edge segment 72 of the slot 70 and may extend to a second end 94 that 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 land segment 74 of the slot 70. In some embodiments, the pressure side damper land 90 may be oriented parallel to the land segment 74 of the slot 70. In other embodiments, the pressure side damper land 90 may be axially oriented relative to the axial centerline of the gas turbine 10.

[0045] The pressure side damar blade 90 can be used to provide a surface for a damar pin 95 to be positioned thereon and to provide vibration damping for the rotor blade 30. In many embodiments, the profile of the surface of the pressure side damar blade 90 can slightly conform to the shape of the damar pin 95 to provide increased surface contact and vibration damping. For example, the pressure side damar blade 90 further includes a curved portion 96 and a flat portion 98. The curved portion 96 can curve circumferentially inward from the platform 42 to the flat portion 98. The flat portion 98 of the pressure side damar blade 90 can extend radially inward from the curved portion 96 to the shank cutout 39 defined in the body 35. The flat portion 98 can be generally parallel to the platform 42 with respect to both the axial and radial directions of the gas turbine 10.

[0046] In some embodiments, the pressure side damar blade 90 can be substantially cantilevered due to the slot 70 and its flat portion 98. For example, the flat portion 98 of the pressure side damar blade 90 can extend radially inward from the curved portion 96 to a free end 99. The free end 99 can be substantially cantilevered within the shank cutout 39 to advantageously provide increased compliance throughout the platform 42 of the rotor blade 30, effectively increasing vibration damping. In various embodiments, the flat portion 98 of the pressure side damar blade can taper axially inward from the curved portion 96 to the free end 99. In various embodiments, the flat portion 98 of the pressure side damar blade 90 can taper away from the first undercut 100 and the second undercut 102 at its respective ends.

[0047] In addition to providing a housing for the seal 84, the slot 70 can provide reduced material stiffness and increased compliance in the pressure side damar blade 90, which allows for increased vibration damping throughout the blade 30. Additionally, the slot 70 can include a slot depth 71. Varying the slot depth (i.e., increasing or decreasing) 71 can advantageously increase or decrease the overall stiffness of the pressure side blade 90, resulting in an increase in overall damping effect.

[0048] In some embodiments, such as Figures 5 to 7 As shown, the pressure side impact surface 56 can further include a first undercut 100 and a second undercut 102. The first undercut 100 and the second undercut 102 are used to advantageously vary, i.e., increase or decrease, the stiffness of the shank 38 to improve overall vibration damping effect. In Figure 5 In the illustrated embodiment, the first undercut 100 and the second undercut 102 can be semicircular cutouts defined circumferentially inward on the body 35 of the rotor blade 30. In some embodiments, the first undercut 100 and the second undercut 102 can be substantially curved or arcuate. In other embodiments, such as Figure 6 and Figure 7 In the illustrated embodiment, the first undercut 100 and the second undercut 102 can include a plurality of semicircular cutouts or trapezoidal cutouts.

[0049] In many embodiments, the first undercut 100 can be disposed directly radially inward from the first end 92 of the pressure side dam blade band 90, and the second undercut 102 can be axially separate from the first undercut 100 and can be disposed directly radially inward from the second end 94 of the pressure side dam blade band 90. In some embodiments, the first undercut 100 and the second undercut 102 can extend generally radially inward from the first end 92 and the second end 94, respectively, of the pressure side dam blade band 90. In many embodiments, both the first undercut 100 and the second undercut 102 can extend radially inward past the free end 99 of the pressure side dam blade band 90.

[0050] The first undercut 100 and the second undercut 102 can each partially define the slot 70. More specifically, the first undercut 100 can partially define the leading edge segment 72 of the slot 70. Likewise, the second undercut 102 can partially define the trailing edge segment 76 of the slot 70. In various embodiments, the first undercut 100 can be disposed axially between the leading edge segment 72 of the slot 70 and the flat portion 98 of the pressure side dam blade band 90. The second undercut 102 can be disposed axially between the flat portion 98 of the pressure side dam blade band 90 and the trailing edge segment 76 of the slot 70.

[0051] The first undercut 100 and the second undercut 102 can each include a maximum undercut depth 106 defined in the circumferential direction. The maximum undercut depth 106 of the first undercut 100 can be the same as or different from the maximum undercut depth 106 of the second undercut 102. Altering the maximum undercut depth 106 of the first undercut and / or the second undercut 102 will advantageously alter (i.e., increase or decrease) the stiffness of the pressure side dam blade band 90, resulting in an increased damping effect. In some embodiments, the maximum undercut depth 106 of the respective undercut 100, 102 can be up to about 1.5 inches. In other embodiments, the maximum undercut depth 106 can be up to about 1 inch. In some embodiments, the maximum undercut depth 106 can be up to about 0.75 inches. In various embodiments, the maximum undercut depth 106 can be up to about 0.5 inches. In other embodiments, the maximum undercut depth 106 can be up to about 0.25 inches.

[0052] Figure 8 A cross-sectional view of a pair of circumferentially adjacent adjacent rotor blades is shown. As shown, when the rotor blades 30', 30" are so positioned, the pressure side dam blade band 90 of the first rotor blade 30' is aligned with the suction side dam blade band 120 of the adjacent second rotor blade 30". As shown, the pressure side dam blade band 90 of the first rotor blade 30' is aligned with the suction side dam blade band 120 of the adjacent second rotor blade 30". Figure 8As shown, damper pins 95 can be disposed along suction side damper blade band 120. In operation, damper pins 95 move in the direction of arrow 130 and contact both pressure side damper blade band 90 and suction side damper blade band 120 to provide vibration damping to adjacent rotor blades 30', 30".

[0053] Further, the slot depth 71' of pressure side impact face 56 can be different than the slot depth 71" of suction side impact face 58. For example, the slot depth 71' of pressure side impact face 56 can be greater than the slot depth 71" of suction side impact face 58, or vice versa. Generally, the sum of the slot depth 71' of pressure side impact face 56, the width of gap 60 in the circumferential direction (shown in FIG. 4), and the slot depth 71" of suction side impact face 58 can be substantially equal to or slightly greater than the width of seal 84. In various embodiments, seal 84 can be smaller than slot 70 and can have room for thermal expansion within slot 70. Additionally, seal 84 can be sized to allow for manufacturing variations thereof. Figure 4

[0054] For example, in many embodiments, the width of seal 84 can be between about 5% and about 30% of the passage width to allow for both manufacturing variations and thermal expansion within slot 70. Figures 2 to 8 The illustrated embodiments allow for the use of both vibration damping pins 95 and seal 84. In various embodiments, blade 30 can include only vibration damping pins 95, only seal 84, or both vibration damping pins 95 and seal 84.

[0055] This written description uses examples to disclose the application, including the best mode, and also to enable any person skilled in the art to practice the application, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the application is defined by the claims, and can 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 have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.​

Claims

1. A rotor blade (30) for a turbine, the rotor blade (30) comprising: a body (35) having a shank (38), a platform (42), and an airfoil (36) extending radially outward from the shank (38), the body (35) including a pressure side impact surface (56) and a suction side impact surface (58); wherein each of the pressure side impact surface (56) and the suction side impact surface (58) includes a damper land (90), (120) and defines a slot (70); and wherein the damper land (90) of the pressure side impact surface (56) is disposed radially inward from the slot (70) of the pressure side impact surface (56), and the damper land (120) of the suction side impact surface (58) is disposed radially inward from the slot (70) of the suction side impact surface (58); The pressure side impact surface (56) further includes one or more undercuts (100), (102) positioned radially inboard of a first end (92) and a second end (94) of the damper land (90) of the pressure side impact surface (56), wherein the one or more undercuts (100), (102) at least partially define the slot (70) of the pressure side impact surface (56).

2. The rotor blade (30) of claim 1 , wherein the damper land (90) of the pressure side impact surface (56) and the damper land (120) of the suction side impact surface (58) each include a first end (92), (122) and a second end (94), (124), the first end (92), (122) being spaced apart from the second end (94), (124) in an axial direction.

3. The rotor blade (30) of claim 2, wherein the suction side impact surface (58) further comprises one or more undercuts (100), (102) positioned radially inboard of the first end (122) and the second end (124) of the damper land (120) of the suction side impact surface (58), wherein the one or more undercuts (100), (102) Each of (102) has a maximum undercut depth (106) defined in the circumferential direction.

4. The rotor blade (30) of any one of claims 2 to 3, wherein each of the one or more undercuts (100), (102) of the pressure side impact surface (56) has a maximum undercut depth (106) defined in a circumferential direction.

5. The rotor blade (30) of claim 4, wherein each of the one or more undercuts (100), (102) is arcuate.

6. The rotor blade (30) of any one of claims 4-5, wherein the maximum undercut depth (106) of each of the one or more undercuts (100), (102) is up to about 1.5 inches.

7. The rotor blade (30) of any one of claims 1 to 6, wherein the slot (70) of the pressure side impact surface (56) and the slot (70) of the suction side impact surface (58) each include a leading edge segment (72), a platform segment (74), and a trailing edge segment (76); and wherein the leading edge segment (72) is defined along the leading edge surface (52), the platform segment (74) is defined along the platform (42), and the trailing edge segment (76) is defined along the trailing edge surface (54).

8. The rotor blade (30) of claim 7, wherein the damper land (90) of the pressure side impact surface (56) is positioned radially inboard of the platform segment (74) of the slot (70) defined in the pressure side impact surface (56); and wherein the damper land (120) of the suction side impact surface (58) is positioned radially inboard of the platform segment (74) of the slot (70) defined in the suction side impact surface (58).

9. A turbine comprising: compressor section (14); Burner section (16); turbine section (18); A plurality of rotor blades (30) disposed 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), a platform (42), and an airfoil (36) extending radially outward from the shank (38), the body (35) including a pressure side impact surface and a suction side impact surface (58); wherein each of the pressure side impact surface (56) and the suction side impact surface (58), respectively, includes a damper land (90), (120) and defines a slot (70); and wherein the damper land (90) of the pressure side impact surface (56) is disposed radially inward from the slot (70) of the pressure side impact surface (56), and the damper land (120) of the suction side impact surface (58) is disposed radially inward from the slot (70) of the suction side impact surface (58); The pressure side impact surface (56) further includes one or more undercuts (100), (102) positioned radially inboard of a first end (92) and a second end (94) of the damper land (90) of the pressure side impact surface (56), wherein the one or more undercuts (100), (102) at least partially define the slot (70) of the pressure side impact surface (56).

10. The turbine of claim 9, wherein the damper land (90) of the pressure side impact surface (56) and the damper land (120) of the suction side impact surface (58) each include a first end (92), (122) and a second end (94), (124), the first end (92), (122) being spaced apart from the second end (94), (124) in an axial direction.

11. The turbine of claim 10, wherein the suction side impact surface (58) further comprises one or more undercuts (100), (102) positioned radially inward of the first end (122) and the second end (124) of the damper land (120) of the suction side impact surface (58), wherein each of the one or more undercuts (100), (102) has a maximum undercut depth (106) defined in a circumferential direction.

12. The turbine according to any one of claims 10 to 11, wherein each of the one or more undercuts (100), (102) of the pressure side impact surface (56) has a maximum undercut depth (106) defined in a circumferential direction.

13. The turbine of claim 12, wherein the one or more undercuts (100), Each of (102) is arcuate.

14. The turbine of any one of claims 12 to 13, 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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