EP3222811A1 - Schwingungsdämpfung in einer gasturbine - Google Patents

Schwingungsdämpfung in einer gasturbine Download PDF

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Publication number
EP3222811A1
EP3222811A1 EP16162147.9A EP16162147A EP3222811A1 EP 3222811 A1 EP3222811 A1 EP 3222811A1 EP 16162147 A EP16162147 A EP 16162147A EP 3222811 A1 EP3222811 A1 EP 3222811A1
Authority
EP
European Patent Office
Prior art keywords
rotor component
rotor
contact surface
damping structure
arrangement according
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.)
Withdrawn
Application number
EP16162147.9A
Other languages
English (en)
French (fr)
Inventor
Omid Lorestani
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.)
Siemens AG
Original Assignee
Siemens AG
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 Siemens AG filed Critical Siemens AG
Priority to EP16162147.9A priority Critical patent/EP3222811A1/de
Priority to PCT/EP2017/053103 priority patent/WO2017162365A1/en
Publication of EP3222811A1 publication Critical patent/EP3222811A1/de
Withdrawn legal-status Critical Current

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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
    • 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/04—Antivibration arrangements
    • 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
    • F01D5/02—Blade-carrying members, e.g. rotors
    • F01D5/10—Anti- vibration means
    • 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
    • F05D2260/00—Function
    • F05D2260/96—Preventing, counteracting or reducing vibration or noise

Definitions

  • the present invention relates to an arrangement for damping vibrations in a gas turbine and to a method for damping vibrations-in a gas turbine and further relates to a gas turbine comprising the arrangement for damping vibrations in a gas turbine.
  • a gas turbine is conventionally used to produce mechanical power or electrical power by burning a fuel.
  • vibrations may be excited in different turbine engine components. The vibrations may lead to a damage of components of the gas turbine and may even lead to a catastrophic rotor failure.
  • US 4,361,213 discloses a vibration damper ring for damping vibration of a structural member of a gas turbine.
  • the damper ring includes an annular recess for receiving an extension, such as a flange, of the structural member and can include radially extending tabs to restrict relative axial movement.
  • the damper ring is preferably severed to allow it to be retained radially against the extension of the structural member.
  • US 5,733,103 discloses a vibration damper for a turbine engine, the damper including a damper ring having inner and outer radial surfaces, axial retainers extending radially outward from the outer radial surface and an annular stiffener protruding radially inward from the inner radial surface.
  • US 6,375,428 B1 discloses a turbine blisk rim friction finger damper for reducing vibrations in an integrally bladed turbine disk.
  • the damper includes an annular member and a plurality of fingers and the annular member is configured so that it is coupled to a face of the integrally bladed turbine disk.
  • an arrangement for damping vibrations in a gas turbine comprising: a rotor arranged for rotating, a first rotor component connected to the rotor, a second rotor component connected to the rotor, and a damping structure fixed at the first rotor component or floated between the two rotor components and having a contact surface arranged to press against a contact surface of the second rotor component to damp a vibration or vibration of the second rotor component.
  • the first rotor component may for example comprise or be a disk at which plural rotor blades are connected.
  • the first rotor component may in particular be a component which does not oscillate or at least oscillates only with much smaller amplitude than the second rotor component or which vibrates within acceptable amplitudes.
  • the second rotor component may in particular be a component which is axially immediately adjacent to the first rotor component.
  • the second rotor component may in particular not comprise any rotor blades connected to it.
  • the second rotor component may for example be used for or be part of a sealing or system preventing mixture of a cooling gas, such as compressed air with hot gas, or it could be used as a balancing piston to cancel out axial forces in the rotor.
  • the second rotor component may be axially spaced apart from the first rotor component.
  • a number of vibrations may occur. The vibrations may be in the axial direction, in the radial direction and/or in the circumferential direction.
  • Embodiments of the present invention are in particular directed for damping or reducing vibrations of the second rotor component which are in the axial direction.
  • the damping structure may also be effective for reducing vibrations in the radial and/or the circumferential direction.
  • the damping structure may be made of an alloy or a super alloy a pure metal and may in particular be configured to have a required rigidity and thermal resistance.
  • the damping structure may be surrounded for example by a cooling gas, which may have a temperature between 300°C and 500°C. Further, during operation, high centrifugal forces may act on the damping structure which the damping structure should withstand.
  • the damping structure may be an annular structure running the circumferential direction and forming a closed ring.
  • the damping structure may by itself be rigid but may slightly elongate due to centrifugal forces and/or thermal expansion.
  • the contact surface of the damping structure via which the damping structure touches and presses against the contact surface of the second rotor component may be spaced apart from a portion where the damping structure is fixed at the first rotor component.
  • second rotor component may move (in particular vibrate) relative to the damping structure and friction may occur between the contact surface of the damping structure and the contact surface of the second rotor component. The friction may reduce the movement of the damping structure and the second rotor component relative to each other and may thus damp the vibration and reduces the amplitude of the vibration of the second rotor component.
  • the damping structure is designed to exert, during rotation of the rotor at a predetermined rotational speed, a predetermined force to the contact surface of the second rotor component, in order to damp an axial vibration of the second rotor component, the contact surface of the second rotor component in particular comprising a coating configured to reduce wear due to friction against the contact surface of the damping surface.
  • the damping structure (in particular the shape of its contact surface, etc.) may be designed such as to exert a required pressing force towards the contact surface of the second rotor component in order to damp vibrations having an expected amplitude.
  • the surface shapes of the contact surface of the damping structure as well as the contact surface of the second rotor component may be determined and the shape and extension of the damping structure as a whole may be designed, in order to achieve a desired damping force on the contact surface of the second rotor component during operation, i.e. at a particular rotational velocity.
  • the coating of the region of the contact surface of the second rotor component may comprise a particular alloy layer in order to increase the durability and thereby reduce wear at this region of the second rotor component.
  • the contact surface of the second rotor component has an angle against an axial direction between 30° and 75°, the contact surface having in particular a curved shape, wherein in particular the predetermined force has a predetermined direction.
  • the pressing force may act substantially orthogonal to the surface and may then effectively damp vibrations of the second rotor component.
  • the direction of the force may be adjusted, to achieve an effective damping.
  • the damping structure can comprises a fixing section used to fix the damping structure at the first rotor component, wherein the fixing section of the damping structure is located further radially inwards than the contact surface of the damping structure.
  • the fixing section may for example comprise a hole to be bolted or hanged, or may comprise a region which may be welded with the first rotor component.
  • a radially outward directed centrifugal force may act on the contact surface of the damping structure and may effectively press the contact surface of the damping structure onto the contact surface of the second rotor component, to effect damping the vibrations of the second rotor component.
  • the damping structure further comprises a spacer portion between the fixing section and the contact surface of the damping structure, the spacer portion forming a ring plate.
  • the spacer portion may allow to protrude to the contact surface of the second rotor component which is best suitable for applying a force for damping the vibrations.
  • the spacer portion may run in the radial and/or in the axial direction. Due to the annular shape of the damping structure, the spacer region also extends in the circumferential direction.
  • the ring plate may be formed by a ring having a constant or variable thickness.
  • the contact surface of the damping structure may have a shape different from the spacer portion in order to smoothly match the surface shape of the contact surface of the second rotor component.
  • the fixing section allows bolting the damping structure to the first rotor component. Thereby, a simple fixing method may be realized.
  • the damping structure extends to a larger degree radially than axially.
  • a centrifugal force acting during operation may effectively increase a pressing force with increasing rotational speed, in order to effectively damp amplitudes of vibrations which may also increase with increasing rotational speed.
  • the damping structure forms a ring, made from a single piece or comprising several sections of a segmented ring.
  • the ring comprises several sections, it may be possible to damp local vibrations, i.e. vibrations which do not occur at all circumferential positions.
  • the ring sections may at least partly be uncoupled from each other although they may be connected to some degree or even rigidly connected in other embodiments.
  • the arrangement is part of a turbine section of the gas turbine, wherein the first rotor component may be a disk having rotor blades connected to it.
  • the disk may in particular be a second, third or fourth or fifth disk of the turbine section, i.e. a disk not immediately downstream of an exhaust region of a combustor.
  • the second rotor component may comprises at a radially most outward section plural teeth of a labyrinth seal formed against a stator portion, the labyrinth seal being intended to inhibit flow through of a cooling gas.
  • the plural teeth together with components of a stator portion may form the labyrinth seal.
  • the cooling gas may for example be compressed air generated in a compressor section of the gas turbine.
  • the cooling gas may have a temperature between 300°C and 500°C. Other values are possible.
  • the contact surface of the damping structure contacts the second rotor component radially inwards from the plural teeth.
  • the seal function may not be affected or in particular not deteriorated by providing the damping structure.
  • flow through of the cooling gas through the labyrinth seal excites vibration of the second rotor component.
  • Other causes of the vibration may be present.
  • the contact portion of the second rotor component is radially inwards from the teeth. Therefore, the sealing function is not affected, while vibrations may effectively be damped.
  • a gas turbine comprising an arrangement for damping vibration according to one of the preceding embodiments.
  • a method for damping vibrations in a gas turbine comprising: rotating a rotor, exciting a vibration of a second rotor component connected to the rotor, pressing against a contact surface of the second rotor component using a contact surface of a damping structure fixed at a first rotor component that is connected to the rotor.
  • Fig. 1 shows an example of a gas turbine engine 10 according to an embodiment of the present invention in a sectional view.
  • the gas turbine engine 10 comprises, in flow series, an inlet 12, a compressor section 14, a combustor section 16 and a turbine section 18 which are generally arranged in flow series and generally about and in the direction of a longitudinal or rotational axis 20.
  • the gas turbine engine 10 further comprises a rotor shaft 22 which is rotatable about the rotational axis 20 and which extends longitudinally through the gas turbine engine 10.
  • the rotor shaft 22 drivingly connects the turbine section 18 to the compressor section 14.
  • air 24, which is taken in through the air inlet 12 is compressed by the compressor section 14 and delivered to the combustion section or burner section 16.
  • the burner section 16 comprises a burner plenum 26, one or more combustion chambers 28 and at least one burner 30 fixed to each combustion chamber 28.
  • the combustion chambers 28 and the burners 30 are located inside the burner plenum 26.
  • the compressed air passing through the compressor 14 enters a diffuser 32 and is discharged from the diffuser 32 into the burner plenum 26 from where a portion of the air enters the burner 30 and is mixed with a gaseous or liquid fuel.
  • the air/fuel mixture is then burned and the combustion gas 34 or working gas from the combustion is channelled through the combustion chamber 28 to the turbine section 18 via a transition duct 17.
  • the exemplary gas turbine engine 10 has an annular combustor section arrangement 16, which is constituted by an annular array of combustor cans 19 each having the burner 30 and the combustion chamber 28, the transition duct 17 has a generally circular inlet that interfaces with the combustor chamber 28 and an outlet in the form of an annular segment.
  • An annular array of transition duct outlets form an annulus for channelling the combustion gases to the turbine 18.
  • the turbine section 18 comprises a number of blade carrying discs 36 attached to the shaft 22. In the present example, two discs 36 each carry an annular array of turbine blades 38. However, the number of blade carrying discs could be different, i.e. only one disc or more than two discs.
  • guiding vanes 40 which are fixed to a stator 42 of the gas turbine engine 10, are disposed between the stages of annular arrays of turbine blades 38. Between the exit of the combustion chamber 28 and the leading turbine blades 38 inlet guiding vanes 44 are provided and turn the flow of working gas onto the turbine blades 38.
  • the combustion gas from the combustion chamber 28 enters the turbine section 18 and drives the turbine blades 38 which in turn rotate the shaft 22.
  • the guiding vanes 40, 44 serve to optimise the angle of the combustion or working gas on the turbine blades 38.
  • the turbine section 18 drives the compressor section 14.
  • the compressor section 14 comprises an axial series of vane stages 46 and rotor blade stages 48.
  • the rotor blade stages 48 comprise a rotor disc supporting an annular array of blades.
  • the compressor section 14 also comprises a casing 50 that surrounds the rotor stages and supports the vane stages 48.
  • the guide vane stages include an annular array of radially extending vanes that are mounted to the casing 50. The vanes are provided to present gas flow at an optimal angle for the blades at a given engine operational point.
  • Some of the guide vane stages have variable vanes, where the angle of the vanes, about their own longitudinal axis, can be adjusted for angle according to air flow characteristics that can occur at different engine operations conditions.
  • the casing 50 defines a radially outer surface 52 of the passage 56 of the compressor 14.
  • a radially inner surface 54 of the passage 56 is at least partly defined by a rotor drum 53 of the rotor which is partly defined by the annular array of blades 48.
  • the present invention is described with reference to the above exemplary turbine engine having a single shaft or spool connecting a single, multi-stage compressor and a single, one or more stage turbine.
  • the present invention is equally applicable to two or three shaft engines and which can be used for industrial, aero or marine applications.
  • upstream and downstream refer to the flow direction of the airflow and/or working gas flow through the engine unless otherwise stated.
  • forward and rearward refer to the general flow of gas through the engine.
  • axial, radial and circumferential are made with reference to the rotational axis 20 of the engine.
  • Fig. 2 illustrates in a sectional view a portion of the turbine section 18 of the gas turbine 10 illustrated in Fig. 1 .
  • the turbine section 18 comprises an arrangement 60 for damping vibrations.
  • the arrangement 60 comprises a rotor shaft 22, a first rotor component 62 connected to the rotor, a second rotor component 64 connected to the rotor and a damping structure 66 fixed at the first rotor component 62 or floated between first rotor component 62 and second one 64 and having a contact surface 68 arranged to press against a contact surface 70 of the second rotor component 64 to damp a vibration of the second rotor component 64.
  • the first rotor component 62 is formed by the disk 36 which carries a rotor blade 38, e.g. a third or fourth stage downstream the combustor or can be a disk that carries no blades.
  • the second rotor component 64 is formed by a disk which does not carry rotor blades but which carries plural teeth 72 which, together with a profile 74 at a stator portion 76 form a labyrinth seal 78 to inhibit flow-through of a cooling gas 80.
  • the damping structure 66 is designed to exert, during rotation of the rotor shaft 22 at a predetermined rotational speed, a predetermined force F to the contact surface 70 of the second rotor component 64.
  • a tangent 69 taken on the contact surface 68 of the damping structure 66 forms an angle ⁇ with the axial direction 82 (The radial direction is labelled with reference sign 83).
  • the angle ⁇ may be between 30° and 75. Other values are possible.
  • the contact surface 68 of the damping structure 66 has a curved shape, in particular being convex, the contact surface 70 of the second component having a concave shape.
  • the contact surface 68 (and also the contact surface 70 of the second component) has a plane or planar shape.
  • the damping structure 66 comprises a fixing section 84 which allows for example to fix the damping structure 66 via a bolt 86 (or some other mechanism) to the first rotor component 62, i.e. the disk 36.
  • the radial position of the fixing section is r1 and the radial position of the (e.g. center of the) contact surface 68 of the damping structure 66 is r2, wherein r2 > r1.
  • the damping structure 66 further comprises a spacer portion 88 between the fixing section 84 and the contact surface 68.
  • the contact surface 68 of the damping structure 66 may advantageously be positioned to be in tight contact and to press against the contact surface 70 of the second rotor component.
  • the damping structure 66 has a radial extent sr and has an axial extent toward the second rotor component 64 sa, wherein sa ⁇ sr.
  • the teeth 72 are radially outward from the contact surface 68 of the damping structure 66.
  • the damping structure consists of a membrane (which also could be divided into segments).
  • the membrane may be mounted (floated, bolted, hooked, etc.) behind the component having the vibrations to be damped.
  • the membrane may be elongated (due to centrifugal force and due to thermal expansion) and the contact with the disk 64 may be retained. Vibration of the disk 64 results in a relative motion between the disk 64 and the membrane 66. Due to a relative motion and friction between contact surfaces, damping will be introduced in the system which helps to reduce disk vibrations. It also adds additional stiffness to the critical location on disk which is favorable.
  • Embodiments of the present invention allow the possibility to solve vibration problems in engines with minimum modifications and changes to the engine. Even it allows to easily solve the problem in machines which are currently operating in the field.
EP16162147.9A 2016-03-24 2016-03-24 Schwingungsdämpfung in einer gasturbine Withdrawn EP3222811A1 (de)

Priority Applications (2)

Application Number Priority Date Filing Date Title
EP16162147.9A EP3222811A1 (de) 2016-03-24 2016-03-24 Schwingungsdämpfung in einer gasturbine
PCT/EP2017/053103 WO2017162365A1 (en) 2016-03-24 2017-02-13 Damping vibrations in a gas turbine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP16162147.9A EP3222811A1 (de) 2016-03-24 2016-03-24 Schwingungsdämpfung in einer gasturbine

Publications (1)

Publication Number Publication Date
EP3222811A1 true EP3222811A1 (de) 2017-09-27

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Family Applications (1)

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EP16162147.9A Withdrawn EP3222811A1 (de) 2016-03-24 2016-03-24 Schwingungsdämpfung in einer gasturbine

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EP (1) EP3222811A1 (de)
WO (1) WO2017162365A1 (de)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN113606000A (zh) * 2021-07-28 2021-11-05 中国科学院工程热物理研究所 一种具有减振减重功能的盘式转子系统
CN114026311A (zh) * 2019-05-29 2022-02-08 赛峰飞机发动机公司 具有阻尼器的涡轮机组件

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN114109529B (zh) * 2021-11-30 2023-12-19 中国航发湖南动力机械研究所 一种弹支挤压油膜金属橡胶阻尼器

Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4361213A (en) 1980-05-22 1982-11-30 General Electric Company Vibration damper ring
US5733103A (en) 1996-12-17 1998-03-31 General Electric Company Vibration damper for a turbine engine
US6375428B1 (en) 2000-08-10 2002-04-23 The Boeing Company Turbine blisk rim friction finger damper
EP1584785A1 (de) * 2004-04-09 2005-10-12 Snecma Kompensationsmasse zur Auswuchtung eines Rotors, besonders für einen Rotor eines Flugtriebwerks
EP1602855A2 (de) * 2004-06-01 2005-12-07 General Electric Company Einrichtung zum Auswuchten eines Turbinenrotors
US20120207603A1 (en) * 2009-06-16 2012-08-16 General Electric Company Trapped spring balance weight and rotor assembly

Family Cites Families (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3666376A (en) * 1971-01-05 1972-05-30 United Aircraft Corp Turbine blade damper
US4101245A (en) * 1976-12-27 1978-07-18 United Technologies Corporation Interblade damper and seal for turbomachinery rotor
US8066479B2 (en) * 2010-04-05 2011-11-29 Pratt & Whitney Rocketdyne, Inc. Non-integral platform and damper for an airfoil

Patent Citations (6)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4361213A (en) 1980-05-22 1982-11-30 General Electric Company Vibration damper ring
US5733103A (en) 1996-12-17 1998-03-31 General Electric Company Vibration damper for a turbine engine
US6375428B1 (en) 2000-08-10 2002-04-23 The Boeing Company Turbine blisk rim friction finger damper
EP1584785A1 (de) * 2004-04-09 2005-10-12 Snecma Kompensationsmasse zur Auswuchtung eines Rotors, besonders für einen Rotor eines Flugtriebwerks
EP1602855A2 (de) * 2004-06-01 2005-12-07 General Electric Company Einrichtung zum Auswuchten eines Turbinenrotors
US20120207603A1 (en) * 2009-06-16 2012-08-16 General Electric Company Trapped spring balance weight and rotor assembly

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
CN114026311A (zh) * 2019-05-29 2022-02-08 赛峰飞机发动机公司 具有阻尼器的涡轮机组件
CN114026311B (zh) * 2019-05-29 2024-04-02 赛峰飞机发动机公司 具有阻尼器的涡轮机组件
CN113606000A (zh) * 2021-07-28 2021-11-05 中国科学院工程热物理研究所 一种具有减振减重功能的盘式转子系统
CN113606000B (zh) * 2021-07-28 2023-03-14 中国科学院工程热物理研究所 一种具有减振减重功能的盘式转子系统

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