Damper stack for turbomachinery rotor blades - Patents.com

Abstract

To provide a damper stack, a rotor blade, and a turbomachine.SOLUTION: A rotor blade comprises a shank, and a body that comprises an airfoil (36) extending radially outward from the shank. The rotor blade further comprises a damping passage (60) defined by the body, where the damping passage extends radially through the body. The rotor blade further comprises a damper stack (70) located in the damping passage, the damper stack comprises a plurality of damper pins (72), and each of the plurality of damper pins contacts an adjacent damper pin.SELECTED DRAWING: Figure 3

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, the damping function of such damper designs is often limited to damping the relative motion between adjacent rotor blades. Furthermore, such damper designs may wear during operation and often become ineffective after a period of time due to such wear.

[0006]

[0005] Accordingly, improved damper designs are desired in the art. In particular, a damper design that provides damping of the absolute motion of a single rotor blade, regardless of the relative motion between adjacent blades, would be advantageous. Additionally, a damper design that continues to provide damping despite wear during operation 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 damping passage defined in the body, the damping passage extending radially through the body. The rotor blade further includes a damper stack disposed within the damping passage, the damper stack including 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, a turbine section, and 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 damping passage defined in the body, the damping passage extending radially through the body. Each of the plurality of rotor blades further includes a damper stack disposed within the damping passage, the damper stack including a plurality of damper pins, each of the plurality of damper pins contacting an adjacent damper pin.

[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] 2 is a cross-sectional view of a portion of a rotor blade taken generally along a tangential direction in accordance with an embodiment of the present disclosure. [Figure 4] 1 is a generally axial cross-sectional view of a portion of a rotor blade according to an embodiment of the present disclosure. [Diagram 5]4 is a generally axial cross-sectional view of a portion of a rotor blade according to another embodiment of the present disclosure. [Figure 6] 4 is a generally axial cross-sectional view of a portion of a rotor blade according to yet another embodiment of the present disclosure. [Figure 7] FIG. 2 is a perspective view of a damper stack according to an embodiment of the present disclosure. [Figure 8] FIG. 2 is a cross-sectional view of a damper stack according to an embodiment of the present disclosure. [Figure 9] FIG. 13 is a cross-sectional view of a damper stack according to another embodiment of the present disclosure. [Figure 10] 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 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.

[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 illustrates one 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 the intersection or transition between the airfoil 36 and the shank 38, as shown, and may extend generally axially and tangentially outward.

[0024] 3 and 4, the rotor blade 30 may further include one or more shrouds. For example, the airfoil 36 may extend radially between a base 44 (at the intersection between the airfoil 36 and the shank 38) and a tip 46. In some embodiments, a tip shroud 50 may be provided at the tip 46 and may extend generally axially and tangentially outwardly from the airfoil 36. In some embodiments, one or more intermediate shrouds 52 may be provided between the tip 46 and the base 44 and may extend generally axially and tangentially outwardly from the airfoil 36.

[0025] One or more cooling passages 54 may be defined in the body, such as the airfoil 36 and the shank 38. Each cooling passage 54 may extend radially through the body, such as through the airfoil 36 (as shown) and / or the shank 38. Additionally, one or more cooling passages 54 may be connected to form a cooling circuit. FIG. 3 illustrates a first cooling circuit 56 and a second cooling circuit 58, each of which includes a plurality of connected cooling passages 54. A cooling medium may be flowed through the cooling passages 54 to cool the body and rotor blades 30 during operation.

[0026] 3-6, one or more damping passages 60 may be defined in and extend radially through a body, such as the airfoil 36 (as shown), as well as the shank 38. In some embodiments, the damping passage 60 may be one of the cooling passages 54. In other embodiments, the damping passage 60 may be separate and independent from the cooling passage 54, such that no coolant flows through the damping passage 60.

[0027] The damping passages 60 may be defined extending radially through the entire body or only a portion thereof. For example, as discussed, at least a portion (which may be the entirety) of the damping passages 60 may be defined extending through the airfoil 36. In some embodiments, as shown in FIG. 4, the portion of the damping passages 60 defined extending through the airfoil 36 may extend from the base 44 through the tip 46. In other embodiments, as shown in FIG. 5, the damping passages 60 extend radially through only a portion of the airfoil 36 and do not extend to the tip 46.

[0028] In some embodiments, a plug 62 may be provided and positioned within the damping passage 60 at the tip 46 or elsewhere within the damping passage 60. The plug 62 may be, for example, brazed, welded, threadedly engaged, or otherwise secured in place at the tip 46 or elsewhere within the damping passage 60. In embodiments in which the damping passage 60 extends through the tip 46, the plug 62 may be provided so that the damping passage 60 is not open outside of the rotor blade 30.

[0029] In some embodiments, as shown in Figures 4 and 5, the damping passage 60 is a single, unobstructed passage. In other embodiments, as shown in Figure 6, the damping passage 60 may be segmented into one or more passage segments. This may allow multiple independent damper stacks 70 (discussed herein) to be utilized in a given damping passage 60. For example, a static insert 64 may be disposed within the damping passage 60. The static insert 64 may be provided in the damping passage 60 during casting or other forming of the rotor blade 30, or may be inserted into the damping passage 60 after forming of the rotor blade 30, such as by drilling an access hole 66 in the rotor blade 30 and inserting the static insert 64 into the damping passage 60 through the access hole 66. In some embodiments, the access hole 66 may be formed at least partially through the intermediate shroud 52.

[0030] 3-10 , one or more damper stacks 70 according to the present disclosure may be provided on the rotor blade 30. Each damper stack 70 may be disposed within the damping passage 60. Each damper stack 70 may include a number of damper pins 72. Each damper pin 72 may contact adjacent damper pins 72 within the damper stack 70 and may further contact the walls defining the damping passage 60.

[0031] 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, by providing such damper stacks 70 internal to individual rotor blades 30, the damper stacks 70 operate to damp absolute motion of the individual rotor blades 30, regardless of relative motion between adjacent blades. Such damper stacks 70, for example, may advantageously damp absolute vibration and / or bending motion of individual rotor blades 30.

[0032] Further, as discussed, a damper stack 70 according to the present disclosure advantageously includes a plurality of damper pins 72 disposed end-to-end in coaxial relationship with the longitudinal axis of the damping passage 60. The primary damping of the damper stack 70 is due to contact between the pins 72 of the damper stack 70. In addition, the damper pins 72 contact the sidewalls that define the damping passage 60 and also provide damping.

[0033] One or more damper stacks 70 may be disposed within the damping passage 60. In some embodiments, only a single damper stack 70 is disposed within the damping passage 60, as shown in Figures 3-5. In other embodiments, multiple damper stacks 70 may be disposed within the damping passage 60, as shown in Figure 6. In the embodiment shown in Figure 6, static inserts 64 are disposed between adjacent damper stacks 70 of the multiple damper stacks 70, thereby dividing adjacent damper stacks 70 into individual passage segments.

[0034] 7-10, 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 longitudinal linear array within the damping passage 60 such that adjacent ends 74, 76 of adjacent damper pins 72 contact one another.

[0035] 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.

[0036] As discussed, contact between adjacent damper pins 72 may provide a primary damping mechanism for damping rotor blades 30. Centrifugal forces during operation cause damper pins 72 to maintain such contact despite wear. Thus, a damper stack 70 according to the present disclosure advantageously continues to provide damping despite wear during operation.

[0037] The ends 74, 76 of adjacent damper pins 72 may have a suitable shape to provide such primary damping. In some embodiments, the adjacent ends 74, 76 of adjacent damper pins 72 may have complementary spherical shapes. For example, as shown in FIGS. 8 and 9, 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, other suitable complementary shapes, such as conical, dome-shaped, etc., that provide suitable damping may be utilized. In other embodiments, the adjacent ends 74, 76 of adjacent damper pins 72 may have mirrored shapes. For example, as shown in FIG. 10, 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 may be utilized provided that the shape provides suitable primary damping.

[0038] 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.

[0039] 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 a damper stack 70 may be the same. For example, as shown in FIGS. 8 and 10, the length 73 of each of the first and second damper pins 72', 72'' may be the same. In other embodiments, as shown in FIG. 9, the length 73 of one or more damper pins 72 in a damper stack 70 may differ from the other damper pins 72 in the stack. For example, as shown in FIG. 9, the length 73 of a first damper pin 72' may be different from the length 73 of a second damper pin 72''.

[0040] As discussed, in some embodiments, the damping passage 60 is a cooling passage 54 (i.e., a passage in fluid communication with a source of cooling medium, such as compressed air). Thus, in the illustrated embodiment, each of the multiple damper pins 72 has a hollow cross-sectional profile, such that an internal passage 78 is defined through the damper pins 72 and the damper stack 70. In these embodiments, the cooling medium may flow through and / or around the damper stack 70. However, it should be understood that the use of damper pins 72 and damper stacks 70 having hollow cross-sectional profiles is not limited to embodiments in which the damping passage 60 is a cooling passage 54. In other embodiments, such damper pins 72 and damper stacks 70 may be utilized in damping passages 60 that are separate and independent from the cooling passage 54. Additionally, in other embodiments, damper pins 72 and damper stacks 70 according to the present disclosure may be solid such that no internal passages are defined therethrough, and these solid damper pins 72 and damper stacks 70 may be utilized in damping passages 60 that are or are not cooling passages 54.

[0041] In some embodiments, a wire (not shown) may extend through one or more damper pins 72 of the damper stack 70. For example, the wire may extend through the interior passage 78 or through a separately defined interior passage, thus leaving the passage 78 empty. The wire 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.

[0042] 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]

[0043] 10. Gas Turbine 12 Entrance Section 14 Compressor Section 16 Combustor section, Combustion section 18 Turbine Section 20 Exhaust Section 22 Shaft 24 rotor disk 26 Rotor blades, compressor blades, buckets 28 Rotor disc 30 Rotor blades, turbine blades, buckets 31 Outer casing 32 Hot Gas Path 34 Combustion Gas 36 Airfoil 38 Shank 40 Base, dovetail 42 Platform 44 Base 46 Tip 50 Tip shroud 52 Intermediate shroud 54 Cooling passage 56 First cooling circuit 58 Second Cooling Circuit 60 Damping Passage 62 Plug 64 Static Inserts 66 Access hole 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

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 damping passage (60) defined in the body, the damping passage (60) extending radially through the body; a damper stack (70) disposed within the damping passage (60), 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); It is equipped with the plurality of damper pins (72) comprising a first damper pin (72') and a second damper pin (72''), each of the plurality of damper pins (72) extending between a first end (74) and a second end (76), the first end (74) of the first damper pin (72') contacting the second end (76) of the second damper pin (72''), the first end (74) of the first damper pin (72') having an outwardly spherical shape and the second end (76) of the second damper pin (72'') having an inwardly 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 each of the plurality of damper pins (72) has a hollow cross-sectional profile.

4. The turbomachine of any one of claims 1 to 3, wherein the damper stack (70) is a plurality of damper stacks (70), each of the plurality of damper stacks (70) being disposed within the damping passage (60).

5. The turbomachine of claim 4 , further comprising a static insert (64) disposed in the damping passage (60) between adjacent damper stacks (70) of the plurality of damper stacks (70).

6. The turbomachine of any of the preceding claims, wherein the damping passage (60) extends radially through only a portion of the airfoil (36).

7. 7. The turbomachine of claim 1, wherein the airfoil (36) extends radially between a base (44) and a tip (46), the damping passage (60) extends through the tip (46), and a plug (62) is disposed within the damping passage (60) at the tip (46).

8. The turbomachine of any preceding claim, wherein the plurality of rotor blades (30) are disposed within 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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