US10385701B2 - Damper pin for a turbine blade - Google Patents

Damper pin for a turbine blade Download PDF

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Publication number
US10385701B2
US10385701B2 US14/844,317 US201514844317A US10385701B2 US 10385701 B2 US10385701 B2 US 10385701B2 US 201514844317 A US201514844317 A US 201514844317A US 10385701 B2 US10385701 B2 US 10385701B2
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US
United States
Prior art keywords
end portion
spring member
damper pin
spring
pin
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Active, expires
Application number
US14/844,317
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English (en)
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US20170067348A1 (en
Inventor
Spencer A. Kareff
Brian Denver Potter
Ariel Caesar Prepena Jacala
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
GE Infrastructure Technology LLC
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General Electric Co
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by General Electric Co filed Critical General Electric Co
Priority to US14/844,317 priority Critical patent/US10385701B2/en
Assigned to GENERAL ELECTRIC COMPANY reassignment GENERAL ELECTRIC COMPANY ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: KAREFF, Spencer A., JACALA, ARIEL CAESAR PREPENA, Potter, Brian Denver
Priority to EP16184881.7A priority patent/EP3139000B1/fr
Priority to JP2016162311A priority patent/JP6827736B2/ja
Priority to CN201610799886.XA priority patent/CN106499445B/zh
Publication of US20170067348A1 publication Critical patent/US20170067348A1/en
Application granted granted Critical
Publication of US10385701B2 publication Critical patent/US10385701B2/en
Assigned to GE INFRASTRUCTURE TECHNOLOGY LLC reassignment GE INFRASTRUCTURE TECHNOLOGY LLC ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: GENERAL ELECTRIC COMPANY
Active legal-status Critical Current
Adjusted expiration legal-status Critical

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/22—Blade-to-blade connections, e.g. for damping vibrations
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/26—Antivibration means not restricted to blade form or construction or to blade-to-blade connections or to the use of particular materials
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/32—Application in turbines in gas turbines
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/25—Three-dimensional helical
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
    • Y02T50/671—
    • Y02T50/673—

Definitions

  • the present invention generally relates to a turbomachine having multiple circumferentially aligned turbine blades. More particularly, this invention involves a damper pin having a spring member for providing vibration damping between adjacent turbine blades.
  • a turbine blade also known as a turbine bucket or turbine rotor blade, converts energy from a flowing fluid such as hot combustion gas or steam into mechanical energy by causing a rotor shaft of a turbomachine to rotate. As the turbomachine transitions through various operating modes, the turbine blades are subjected to both mechanical and thermal stresses.
  • a turbine blade generally includes an airfoil that extends radially outwardly from a platform, a shank that extends radially inwardly from the platform and a dovetail or mounting portion that extends radially inwardly from the shank.
  • the dovetail of each turbine blade is secured within a complementary slot defined in a rotor wheel or disk.
  • the rotor wheel is coupled to the rotor shaft.
  • vibrations may be introduced into the turbine blades. For example, fluctuations in flow of the hot combustion gases or steam may cause them to vibrate.
  • One basic design consideration for turbomachine designers is to avoid or to minimize resonance with natural frequencies of the turbine blades and the dynamic stresses produced by forced response and/or aero-elastic instabilities, thus controlling high cycle fatigue of the turbine blades.
  • vibration dampers are typically provided below and/or between the platforms to frictionally dissipate vibratory energy and reduce the corresponding amplitude of vibration during operation.
  • the amount of vibrational energy that is removed by the vibration damper is a function of the dynamic weight of the vibration damper and the reaction loads.
  • dampers may be largely adequate during typical operations, there is a desire to improve overall damper effectiveness.
  • Prior attempts to accomplish damping of vibrations have included round damper pins, sheet metal flat dampers, or complex wedge shaped dampers. Often true damper performance of these types of dampers is not known until the first engine test. However, at that time, the damper pocket geometry in the turbine blades is locked in by hard tooling. Thus, if the damper does not perform as expected, then a potentially expensive tooling rework may be required. Accordingly, there is desire for a damping pin that provides a natural frequency tuning tool for resonant mode excitation avoidance and that enables independent mode tuning options without necessitating changes to the design of an existing turbine blade.
  • the damper pin for damping adjacent turbine blades coupled to a rotor shaft.
  • the damper pin includes a first end portion that is axially spaced from a second end portion and a spring member that extends axially from an inner surface of the first end portion to an inner surface of the second end portion.
  • the first end portion, the spring member and the second end portion define a generally arcuate top portion of the damper pin.
  • the top portion is configured to contact with a groove defined between the adjacent turbine blades.
  • the turbine engine includes a rotor shaft that extends axially within the turbine engine and an adjacent pair of turbine blades that are coupled to the rotor shaft. Each turbine blade at least partially defines a groove that extends along a slash face of the corresponding turbine blade.
  • the turbine engine further includes a damper pin that is disposed within the groove between the adjacent turbine blades.
  • the damper pin comprises a first end portion that is axially spaced from a second end portion and a spring member that extends axially from an inner surface of the first end portion to an inner surface of the second end portion.
  • the first end portion, the spring member and the second end portion define a generally arcuate top portion of the damper pin.
  • the top portion is configured to contact with the groove defined between the adjacent turbine blades.
  • FIG. 1 illustrates a functional diagram of an exemplary gas turbine as may incorporate at least one embodiment of the present invention
  • FIG. 2 is a perspective view of an exemplary turbine blade according to at least one embodiment of the present invention.
  • FIG. 3 is a schematic illustration of a damper pin disposed between circumferentially adjacent turbine blades according to at least one embodiment of the present invention
  • FIG. 4 is a side view of an exemplary damper pin according to one embodiment of the present invention.
  • FIG. 5 is a top view of the exemplary damper pin as shown in FIG. 4 ;
  • FIG. 6 is a cross sectioned side view of an exemplary damper pin according to one embodiment of the present invention.
  • upstream and downstream refer to the relative direction with respect to fluid flow in a fluid pathway.
  • upstream refers to the direction from which the fluid flows
  • downstream refers to the direction to which the fluid flows.
  • radially refers to the relative direction that is substantially perpendicular to an axial centerline of a particular component
  • axially refers to the relative direction that is substantially parallel and/or coaxially aligned to an axial centerline of a particular component.
  • FIG. 1 illustrates a schematic diagram of one embodiment of a gas turbine 10 .
  • 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) within a combustor section 16 disposed 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 .
  • the gas turbine 10 may include one or more shafts 22 coupled between the compressor section 14 and the turbine section 18 .
  • the turbine section 18 may generally include a rotor shaft 24 having a plurality of rotor disks 26 (one of which is shown) and a plurality of rotor blades 28 extending radially outwardly from and being interconnected to the rotor disk 26 . Each rotor disk 26 in turn, may be coupled to a portion of the rotor shaft 24 that extends through the turbine section 18 .
  • the turbine section 18 further includes an outer casing 30 that circumferentially surrounds the rotor shaft 24 and the rotor blades 28 , thereby at least partially defining a hot gas path 32 through the turbine section 18 .
  • a working fluid such as air flows through the inlet section 12 and into the compressor section 14 where the air is progressively compressed, thus providing pressurized air to the combustors of the combustion section 16 .
  • the pressurized air is mixed with fuel and burned within each combustor to produce combustion gases 34 .
  • the combustion gases 34 flow through the hot gas path 32 from the combustor section 16 into the turbine section 18 , wherein energy (kinetic and/or thermal) is transferred from the combustion gases 34 to the rotor blades 28 , thus causing the rotor shaft 24 to rotate.
  • the mechanical rotational energy may then be used to power the compressor section 14 and/or to generate electricity.
  • the combustion gases 34 exiting the turbine section 18 may then be exhausted from the gas turbine 10 via the exhaust section 20 .
  • FIG. 2 illustrates a conventional turbine blade or bucket 28 including an airfoil 36 , a platform 38 , a shank 40 and a dovetail or mounting portion 42 .
  • FIG. 3 provides a downstream view of a pair of circumferentially adjacent turbine blades 28 ( a ), 28 ( b ).
  • the dovetail 42 is utilized to secure the turbine blade 28 to a periphery of the rotor disk 26 ( FIG. 1 ), as is well understood in the art.
  • the platform 38 defines an inward flow boundary for the combustion gases 34 flowing through the hot gas path 32 of the turbine section 18 ( FIG. 1 ).
  • a damper pin 44 is located along one axial edge (or slash face) 46 adjacent to (i.e., radially inward of) the turbine blade platform 38 . It will be appreciated that a similar damper pin 44 is located between each adjacent pair of turbine blades 28 ( a ), 28 ( b ) ( FIG. 3 ) on the rotor disk 26 ( FIG. 1 ) as apparent from FIG. 3 . In particular embodiments, as shown in FIG. 2 , the damper pin 44 is located in an elongated groove 48 ( FIG. 1 ) that extends along the entire slash face 46 of the turbine blade 28 .
  • the damper pin 44 serves as a vibration damper. When installed, as shown in FIG. 3 , the damper pin 44 is positioned between the adjacent turbine blades 28 ( a ), 28 ( b ). In operation, the damper pin 44 frictionally dissipates vibratory energy and reduces corresponding amplitude of vibration. The amount of vibrational energy that is removed by the damper pin 44 is a function several factors including but not limited to the dynamic weight of the damper pin 44 , the geometry of the damper pin 44 and the reaction loads between the adjacent turbine blades 28 ( a ), 28 ( b ).
  • FIG. 4 provides a side view of an exemplary damper pin 100 according to one embodiment of the present invention.
  • FIG. 5 provides a top view of the damper pin 100 as shown in FIG. 4 . It is to be understood that damper pin 100 shown in FIG. 4 may be substituted for damper pin 44 as shown in FIGS. 2 and 3 .
  • the damper pin 100 includes a first end portion 102 axially spaced from a second end portion 104 with respect to an axial centerline 106 of the damper pin 100 .
  • the first end portion 102 and the second end portion 104 may be coaxially aligned with respect to centerline 106 .
  • the damper pin 100 further includes a spring member 108 that extends axially from an inner surface 110 of the first end portion 102 to an inner surface 112 of the second end portion 104 .
  • the first end portion 102 , the spring member 108 and the second end portion 104 define a generally arcuate top portion or surface 114 of the damper pin 100 .
  • the top portion 114 is generally configured (shaped and/or sized) to contact with a portion of the groove 48 defined between the adjacent turbine blades 28 ( a ), 28 ( b ).
  • the first end portion 102 and/or the second end portion 104 of the damper pin 100 are substantially semi-cylindrical.
  • the first end portion 102 and/or the second end portion 104 may include shoulders 116 , 118 respectfully. This configuration creates flat support surfaces 120 , 122 that are adapted to rest on machined turbine blade platform surfaces or shoulders at opposite ends of the groove 48 formed in the turbine blade slash face 46 , thereby providing support for the damper pin 100 while preventing undesirable excessive rotation during machine operation.
  • opposing ends 124 , 126 of the spring member 108 may be fixedly connected to the first end portion 102 and the second end portion 104 respectfully.
  • the opposing ends 124 , 126 of the spring member 108 may be engaged with or compressed against the inner surface 110 of the first end portion 102 and/or the inner surface 112 of the second end portion 104 .
  • the spring member 108 is generally helical shaped.
  • the spring member is illustrated in the figures as a helical or coil type spring, it is to be understood by one skilled in the art that the spring member 108 may be any suitable type spring such as but not limited to a wave spring or the like and that the invention is not limited to a helical or coil type spring member unless otherwise provided in the claims.
  • the spring member 108 may comprise of multiple springs coaxially aligned and extending between the first end portion 102 and the second end portion 104 .
  • the spring member 108 comprises a first spring 128 coaxially aligned with a second spring 130 .
  • the first spring 128 may be connected at one end 132 to the first end portion 102 and the second spring 130 may be connected at one end 134 to the second end portion 104 .
  • the first and second springs 128 , 130 may be engaged at contact point 136 that is defined between the inner surface 110 of the first end portion 102 and the inner surface 112 of the second end portion 104 .
  • FIG. 6 is a cross sectional side view of an exemplary embodiment of the damper pin 100 according to one embodiment of the present invention.
  • the damper pin 100 may include a retention pin 138 .
  • the retention pin 138 may be coaxially aligned with and disposed between the first end portion 102 and the second end portion 104 .
  • the spring member 108 extends circumferentially around the retention pin 138 .
  • the retention pin 138 may be seated within openings 140 ( a ), 140 ( b ) defined by the first end portion 102 and the second end portion 104 respectfully.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
US14/844,317 2015-09-03 2015-09-03 Damper pin for a turbine blade Active 2038-05-07 US10385701B2 (en)

Priority Applications (4)

Application Number Priority Date Filing Date Title
US14/844,317 US10385701B2 (en) 2015-09-03 2015-09-03 Damper pin for a turbine blade
EP16184881.7A EP3139000B1 (fr) 2015-09-03 2016-08-19 Goupille amortissante pour amortir des aubes de turbine adjacentes couplées à un arbre de rotor et moteur à turbine
JP2016162311A JP6827736B2 (ja) 2015-09-03 2016-08-23 タービンブレード用ダンパピン
CN201610799886.XA CN106499445B (zh) 2015-09-03 2016-08-31 用于涡轮叶片的减振销

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US14/844,317 US10385701B2 (en) 2015-09-03 2015-09-03 Damper pin for a turbine blade

Publications (2)

Publication Number Publication Date
US20170067348A1 US20170067348A1 (en) 2017-03-09
US10385701B2 true US10385701B2 (en) 2019-08-20

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US14/844,317 Active 2038-05-07 US10385701B2 (en) 2015-09-03 2015-09-03 Damper pin for a turbine blade

Country Status (4)

Country Link
US (1) US10385701B2 (fr)
EP (1) EP3139000B1 (fr)
JP (1) JP6827736B2 (fr)
CN (1) CN106499445B (fr)

Cited By (3)

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US20210172326A1 (en) * 2019-12-10 2021-06-10 General Electric Company Damper stacks for turbomachine rotor blades
US20210172325A1 (en) * 2019-12-10 2021-06-10 General Electric Company Damper stacks for turbomachine rotor blades
US11208903B1 (en) 2020-11-20 2021-12-28 Solar Turbines Incorporated Stiffness coupling and vibration damping for turbine blade shroud

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US10648347B2 (en) * 2017-01-03 2020-05-12 General Electric Company Damping inserts and methods for shrouded turbine blades
JP7039355B2 (ja) * 2018-03-28 2022-03-22 三菱重工業株式会社 回転機械
US11242756B2 (en) 2020-05-04 2022-02-08 General Electric Company Damping coating with a constraint layer
US11333026B2 (en) 2020-05-26 2022-05-17 General Electric Company Vibration-damping system for turbomachine blade(s) on spacer adjacent blade stage
CN111677589B (zh) * 2020-06-10 2024-07-19 中国船舶重工集团公司第七0三研究所 复合弹性悬臂式燃气轮机涡轮支撑环减振抗冲组件
US11085303B1 (en) 2020-06-16 2021-08-10 General Electric Company Pressurized damping fluid injection for damping turbine blade vibration
US11143036B1 (en) 2020-08-20 2021-10-12 General Electric Company Turbine blade with friction and impact vibration damping elements

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EP3093439A1 (fr) 2015-05-14 2016-11-16 General Electric Company Système d'amortisseur

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CN106499445A (zh) 2017-03-15
EP3139000A1 (fr) 2017-03-08
US20170067348A1 (en) 2017-03-09
EP3139000B1 (fr) 2023-12-13
JP2017048788A (ja) 2017-03-09
CN106499445B (zh) 2020-04-24

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