EP1717419A1 - Méthode et dispositif pour l'adjustement d'un jeu radial dans un turbomachine et un compresseur axiaux - Google Patents

Méthode et dispositif pour l'adjustement d'un jeu radial dans un turbomachine et un compresseur axiaux Download PDF

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Publication number
EP1717419A1
EP1717419A1 EP05009380A EP05009380A EP1717419A1 EP 1717419 A1 EP1717419 A1 EP 1717419A1 EP 05009380 A EP05009380 A EP 05009380A EP 05009380 A EP05009380 A EP 05009380A EP 1717419 A1 EP1717419 A1 EP 1717419A1
Authority
EP
European Patent Office
Prior art keywords
guide ring
guide
coolant
turbomachine
compressor
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.)
Granted
Application number
EP05009380A
Other languages
German (de)
English (en)
Other versions
EP1717419B1 (fr
Inventor
Tobias Dr. Buchal
Gerhard Hülsemann
Mirko Milazar
Dieter Minninger
Michael Neubauer
Harald Nimptsch
Heinrich Pütz
Kang Dr. Qian
Arnd Dr. Reichert
Volker Dr. Vosberg
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
Priority to EP05009380A priority Critical patent/EP1717419B1/fr
Application filed by Siemens AG filed Critical Siemens AG
Priority to AT05009380T priority patent/ATE484652T1/de
Priority to DE502005010381T priority patent/DE502005010381D1/de
Priority to CN201010175874A priority patent/CN101825003A/zh
Priority to CN2006100752846A priority patent/CN1854468B/zh
Priority to JP2006120073A priority patent/JP2006307853A/ja
Priority to US11/413,871 priority patent/US7766611B2/en
Publication of EP1717419A1 publication Critical patent/EP1717419A1/fr
Application granted granted Critical
Publication of EP1717419B1 publication Critical patent/EP1717419B1/fr
Not-in-force legal-status Critical Current
Anticipated 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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/08—Preventing or minimising internal leakage of working-fluid, e.g. between stages for sealing space between rotor blade tips and stator
    • F01D11/14—Adjusting or regulating tip-clearance, i.e. distance between rotor-blade tips and stator casing
    • F01D11/20—Actively adjusting tip-clearance
    • F01D11/24—Actively adjusting tip-clearance by selectively cooling-heating stator or rotor components
    • 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
    • F01D19/00—Starting of machines or engines; Regulating, controlling, or safety means in connection therewith
    • F01D19/02—Starting of machines or engines; Regulating, controlling, or safety means in connection therewith dependent on temperature of component parts, e.g. of turbine-casing

Definitions

  • the invention relates to a method for adjusting a radial gap formed between a squealer edge of a blade profile and an opposing guide surface of an axial flow-through turbomachine, wherein a guide ring forming the guide surface is acted upon by a coolant. Furthermore, the invention relates to a compressor.
  • the design-related radial gaps are formed between the rotatable blades of the rotor of the turbomachine and the non-rotatably on the stator opposite guide surfaces.
  • the guide surfaces serve to guide the working medium and are formed by circumferentially divided ring segments, which extend coaxially as a guide ring about the axis of rotation of the rotor in the axial direction. During operation of the gas turbine, the blades of the rotor move at a distance from the guide surfaces.
  • freestanding guide vanes may also form a radial gap with respect to a rotating conical or cylindrical guide surface arranged on the rotor.
  • the radial gaps are to be made as small as possible. It is known from the aforementioned Offenlegungsschrift to fasten guide rings to a stator by obliquely arranged holding partners to the radial direction and to displace them during operation of the gas turbine due to the thermal material expansion of the guide ring in the direction of the blade ends towards the reduction of the radial gap.
  • the gap dimension determining design parameters for the warm start of a gas turbine are designed to be as small as possible operating gap, i. Radial gap to meet.
  • the housing cools comparatively quickly with respect to the rotor of the gas turbine.
  • the housing or the guide rings shrink due to their cooling back to their original design size, the initially warm rotor remains initially extended due to the heat stored in it and retarded cools and shrinks.
  • the result is the so-called constricting effect.
  • This situation can cause the radial gap to decrease and the blades of the rotor to touch or even streak the housing or guide ring, which can permanently increase the radial gap or even damage blades.
  • An increased radial gap leads to increased fuel consumption, damaged blades may require early maintenance with corresponding additional costs.
  • Object of the present invention is to provide a method of the type mentioned, which improves the warm start behavior of the turbomachine to increase the availability and at the same time to increase the efficiency.
  • the object of the method is solved by the features of claim 1 and the object directed to the compressor by the features of claim 7.
  • the solution suggests that coolant is applied to the guide ring before starting the turbomachine.
  • the invention is based on the recognition that the hot start conditions of the turbomachine are improved by influencing the radial gap by the gap size of the radial gaps is increased by the proposed method in a still warm or heated, but not in operation turbomachine, compared with the Gap of the radial gap of a gas turbine known from the prior art in the identical state.
  • the hammer-shaped in cross-section guide ring is formed over the circumference by adjacent ring segments. Since the guide ring facing the blades (or vanes) is radially further outward (or inboard), its application of coolant will result in a displacement of the guide surfaces directed away from the opposite squealer edges of the vanes.
  • the increase of the radial gap achieved in this way results in a reduction of the described constricting effect and the risk of being scuffed, which significantly improves the warm start behavior of the turbomachine, ie the turbomachine could be started earlier, relative to its previous departure time.
  • the radial gaps no longer need to be dimensioned after the warm start as unfavorable start of operation.
  • the cooling of the warm guide rings increases the radial gap of the non-operating turbomachine.
  • the radial gap increase obtained for this condition can also be partially utilized, rather than improving the warm start, to reduce the radial gaps of one in an idle and cold condition, i. a smaller lying at ambient temperature fluid flow machine, based on a known from the prior art turbomachine.
  • both the rotor and the housing of the continuous-duty turbomachine warm up to a maximum operating temperature. Both housing and rotor expand, so that the risk of constriction no longer exists. Accordingly, the method is particularly advantageous if the admission of the guide ring with coolant is set during the start of the turbomachine. After reaching the maximum operating temperature, the temperature-induced strains of the turbomachine, ie the stator and the Rotor, completed. Consequently, the guide ring also heats up so that it expands and shifts its guide surface in the direction of the scrape edges of the blades, which leads to an efficiency-increasing reduction of the radial gaps. In particular, when the turbomachine is designed as a compressor of a gas turbine, in which the guide rings are usually uncooled during operation, this can be used advantageously.
  • coolant is taken from an external coolant source.
  • coolant in the form of cooling air is usually taken from the compressor. Since the method is used before the start of the gas turbine, this is not possible.
  • an external coolant source such as a separately driven auxiliary compressor or external fan, must be used to provide the coolant for cooling the guide rings prior to the warm start of the gas turbine.
  • Preferred dimensions can be acted upon by the start of the turbomachine of the guide ring with a heating medium.
  • the turbomachine is, for example, a compressor or a turbine of a gas turbine and the methods known from the prior art, in which material expansions of the guide ring are used for adjusting the radial gap, is applied to the guide ring of a compressor.
  • air or steam can be used as the heating means.
  • FIG. 1 shows, by way of example for a turbomachine, a gas turbine 1 in a longitudinal partial section. It has inside a rotatably mounted about a rotation axis 2 rotor 3, which is also referred to as a turbine runner. Along the rotor 3 follow one another a suction housing 4, a compressor 5, a toroidal annular combustion chamber 6 with a plurality of coaxially arranged burners 7, a turbine unit 8 and the exhaust housing 9.
  • the annular combustion chamber 6 forms a combustion chamber 17 which communicates with an annular flow channel 18.
  • There four successive turbine stages 10 form the turbine unit 8. Each turbine stage 10 and each compressor stage is formed of two blade rings.
  • a compressor stage is formed by a blade row 13 with a ring of guide blade 12 following in the flow direction of the air to be compressed.
  • the blade 15 is radially outside a guide ring 21 and the guide vane 12 radially inward of a guide ring 23 against.
  • the guide rings 21, 23 delimit the flow channel 18 extending in the axial direction of the rotor 3 in the radial direction.
  • the guide rings 21, 23 may be formed from over the circumference adjacent ring segments.
  • FIG. 2 shows the detail II from FIG. 1, a cross-section through a guide ring 21 with an opposite blade, after all temperature-induced strains have been completed.
  • the device shown in FIG. 2 can be provided both in the turbine unit 8 and / or in the compressor 5 of the gas turbine 1.
  • the blades each have a blade profile 19, which is drop-shaped in cross-section, and has a front edge 20, which can be flowed on by a working medium, and a rear edge 22.
  • a cylindrical or conically extending to the axis of rotation 2 of the gas turbine rotor 3 wall 25 forms part of a rotationally fixed inner housing 27.
  • the wall 25 encloses the annular flow channel 18.
  • In the inner housing 27 and in the Wall 25 is a running in the circumferential direction and in cross-section hammer-shaped groove 29 incorporated, in which the guide ring 21 is arranged.
  • the guide ring 21 surrounds the flow channel 18 coaxial with the axis of rotation 2 of the rotor.
  • an insulating layer 26 may be formed, which shields the guide ring 21 thermally against the wall 25 and insulated so that the wall 25 and the inner housing 27 does not shrink also in the direction of the blade.
  • the guide ring 21 is made of a material which under the action of heat, i. a temperature increase, expands, preferably expands more than the wall 25 and the inner housing 27, i. the guide ring 21 has a greater coefficient of thermal expansion than the wall 25 and the inner housing 27th
  • the guide ring 21 is formed substantially corresponding to the hammer-shaped groove 29 and is directly on the back, or as shown, on the insulating layer 26 on the groove bottom of the groove 29 and the front to a contact surface 50 of the undercut 31, so that the guide ring 21 is fixed.
  • the contact surface 50 determines the radial position of the guide ring 21 and is radially further out (or inside) arranged as the tips of the blades 15 (and vanes 12) opposite guide surface 33rd
  • the flow channel 18 facing the guide surface 33 of the guide ring 21 is the blade 15, in particular their squeal edge 35 opposite. Between the squeal edge 35 of each blade 15 and the guide surface 33, a radial gap 36 is formed. During operation of the gas turbine, the blade rotates 15 under the surface 33 away, this is to illustrate the axis of rotation 2 - not true to scale - indicated.
  • a groove 39 provided with the wall 25 or, if present, the insulating layer 26 extending in the circumferential direction, i. form annular supply channel 41.
  • cooling channels 43 which communicate with the supply channel 41 via radial connection channels 45.
  • the housing After switching off the gas turbine 1, the housing cools faster than the rotor 3, so that the expansions of the housing decrease faster or go back and constrict the still warm and thus more extended rotor 3. This reduces the gap of the radial gap 36.
  • coolant 51 is supplied through the supply channel 49 to the supply channel 41, which passes from there via the connection channels 45 in the cooling channels 43 and the guide ring 21 cools.
  • the coolant 51 absorbs the heat still stored in the guide ring 21 and is then blown through openings, not shown, either in the flow channel 18 or returned to the outside via the return channels also not shown from the inside of the machine.
  • the temperature-induced material expansions of the guide ring 21 go back.
  • the guide surface 33 delimiting the flow channel 16 shifts radially outward into the position 33 '.
  • the radial gap 36 increases by the distance X to 36 ', thereby reducing the risk of rubbing the blades 15 on the guide surface 33 or 33' in the case of warm start. This effect can be used to reduce the time between shutdown and warm start of the gas turbine.
  • the method is particularly effective when the guide ring 21 is insulated from the wall 25.
  • only the guide ring 21 is cooled, and not beyond the wall 25. This leads to a particularly efficient cooling of the guide ring 21 and prevents the wall 25 also moves along the same way. This ensures that only the guide ring 21 takes back its heat-related strains.
  • the housing After or during the start, ie during the starting process of the gas turbine 1, the housing heats up and expands. The housing and the inner housing 27 move radially outward. The risk of rubbing the blades 15 with their squeal edge 35 on the guide surface 33 of the guide rings 21 is reduced, so that after a predetermined period of operation, the cooling of the guide rings 21 can be adjusted.
  • the gas turbine 1 continues to heat up until a temperature distribution that is no longer changing has set in it.
  • the material of the guide ring 21 allows a further increase in temperature, even in place of the coolant 51 during operation of the gas turbine 1, a heating medium through the channels 49, 41, 45 are passed.
  • a further increase in temperature of the guide ring 21 causes an additional expansion in the radial direction, with which the radial gap 36 is further reduced. This leads to an increase in efficiency, since less working fluid - in the compressor 5, the gas to be compressed and in the turbine unit 8, the expanding hot gas 11 - can escape unused by the reduced radial gap 36.
  • the radial gap 36 may not only be formed between a radially outer guide surface 33 and a blade 15, but it may also be between the non-rotating guide vane 12 and arranged on the rotor 3 guide surface 23. Accordingly, wall 25 would be part of the rotor 3, so that the guide ring 23 is opposite a guide vane 12. In this case, the directions of displacement also change from outside to inside.
  • the inventive method for changing the radial gaps 36 is particularly suitable for compressor 5. However, it can also be used in the turbine unit 8.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Structures Of Non-Positive Displacement Pumps (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP05009380A 2005-04-28 2005-04-28 Méthode et dispositif pour l'adjustement d'un jeu radial d'un compresseur axial dans une turbomachine Not-in-force EP1717419B1 (fr)

Priority Applications (7)

Application Number Priority Date Filing Date Title
AT05009380T ATE484652T1 (de) 2005-04-28 2005-04-28 Verfahren und vorrichtung zur einstellung eines radialspaltes eines axial durchströmten verdichters einer strömungsmaschine
DE502005010381T DE502005010381D1 (de) 2005-04-28 2005-04-28 Verfahren und Vorrichtung zur Einstellung eines Radialspaltes eines axial durchströmten Verdichters einer Strömungsmaschine
EP05009380A EP1717419B1 (fr) 2005-04-28 2005-04-28 Méthode et dispositif pour l'adjustement d'un jeu radial d'un compresseur axial dans une turbomachine
CN2006100752846A CN1854468B (zh) 2005-04-28 2006-04-18 用于调节轴流式涡轮机和压缩机的径向间隙的方法及装置
CN201010175874A CN101825003A (zh) 2005-04-28 2006-04-18 用于调节轴流式涡轮机和压缩机的径向间隙的方法
JP2006120073A JP2006307853A (ja) 2005-04-28 2006-04-25 軸流流体機械におけるラジアル隙間の調整方法と圧縮機
US11/413,871 US7766611B2 (en) 2005-04-28 2006-04-28 Method for setting a radial gap of an axial-throughflow turbomachine and compressor

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP05009380A EP1717419B1 (fr) 2005-04-28 2005-04-28 Méthode et dispositif pour l'adjustement d'un jeu radial d'un compresseur axial dans une turbomachine

Publications (2)

Publication Number Publication Date
EP1717419A1 true EP1717419A1 (fr) 2006-11-02
EP1717419B1 EP1717419B1 (fr) 2010-10-13

Family

ID=35765672

Family Applications (1)

Application Number Title Priority Date Filing Date
EP05009380A Not-in-force EP1717419B1 (fr) 2005-04-28 2005-04-28 Méthode et dispositif pour l'adjustement d'un jeu radial d'un compresseur axial dans une turbomachine

Country Status (6)

Country Link
US (1) US7766611B2 (fr)
EP (1) EP1717419B1 (fr)
JP (1) JP2006307853A (fr)
CN (2) CN101825003A (fr)
AT (1) ATE484652T1 (fr)
DE (1) DE502005010381D1 (fr)

Cited By (8)

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Publication number Priority date Publication date Assignee Title
DE102008005482A1 (de) * 2008-01-23 2009-07-30 Rolls-Royce Deutschland Ltd & Co Kg Gasturbine mit einem Verdichter mit selbstheilender Einlaufschicht
EP2351912A1 (fr) * 2010-01-12 2011-08-03 Siemens Aktiengesellschaft Turbine avec système chauffant, centrale énergétique solaire et procédé d'exploitation associés
US8257016B2 (en) 2008-01-23 2012-09-04 Rolls-Royce Deutschland Ltd & Co Kg Gas turbine with a compressor with self-healing abradable coating
EP2273073A3 (fr) * 2009-06-12 2013-07-03 Rolls-Royce plc Système et procédé pour régler le jeu rotor-stator
FR3002273A1 (fr) * 2013-02-20 2014-08-22 Snecma Dispositif avionique pour la surveillance d'une turbomachine
EP2955327A1 (fr) * 2014-06-10 2015-12-16 Rolls-Royce plc Un ensemble d'étanchéité de pointe d'aube
FR3045717A1 (fr) * 2015-12-22 2017-06-23 Snecma Dispositif de pilotage de jeu en sommets d'aubes rotatives de turbine
EP3354859A1 (fr) * 2017-01-26 2018-08-01 Safran Aero Boosters SA Système de controle actif de jeu radial de turbomachine et turbomachine associée

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US8152457B2 (en) * 2009-01-15 2012-04-10 General Electric Company Compressor clearance control system using bearing oil waste heat
US8177476B2 (en) * 2009-03-25 2012-05-15 General Electric Company Method and apparatus for clearance control
EP2282014A1 (fr) * 2009-06-23 2011-02-09 Siemens Aktiengesellschaft Section de canal d'écoulement annulaire pour une turbomachine
RU2012132193A (ru) * 2009-12-30 2014-02-10 Сименс Акциенгезелльшафт Турбина для преобразования энергии и способ ее работы
GB201021327D0 (en) 2010-12-16 2011-01-26 Rolls Royce Plc Clearance control arrangement
US9057282B2 (en) * 2011-11-22 2015-06-16 General Electric Company Systems and methods for adjusting clearances in turbines
JP6643225B2 (ja) 2013-06-11 2020-02-12 ゼネラル・エレクトリック・カンパニイ クリアランス制御リング組立体
US20150047358A1 (en) * 2013-08-14 2015-02-19 General Electric Company Inner barrel member with integrated diffuser for a gas turbomachine
US11732892B2 (en) 2013-08-14 2023-08-22 General Electric Company Gas turbomachine diffuser assembly with radial flow splitters
ITFI20130237A1 (it) * 2013-10-14 2015-04-15 Nuovo Pignone Srl "sealing clearance control in turbomachines"
US20160326915A1 (en) * 2015-05-08 2016-11-10 General Electric Company System and method for waste heat powered active clearance control
CN104963729A (zh) * 2015-07-09 2015-10-07 中国航空工业集团公司沈阳发动机设计研究所 重型燃气轮机高涡叶尖间隙控制结构
US10738791B2 (en) 2015-12-16 2020-08-11 General Electric Company Active high pressure compressor clearance control
US20170306775A1 (en) * 2016-04-21 2017-10-26 General Electric Company Article, component, and method of making a component
US10724535B2 (en) * 2017-11-14 2020-07-28 Raytheon Technologies Corporation Fan assembly of a gas turbine engine with a tip shroud
CN112922904B (zh) * 2021-03-03 2022-10-11 西华大学 基于中介机匣导流的压气机新型扩稳结构
CN116557349B (zh) * 2023-05-18 2024-05-17 中国船舶集团有限公司第七〇三研究所 一种双层交错式压气机机匣处理结构

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EP0559420A1 (fr) * 1992-03-06 1993-09-08 General Electric Company Virole de réglage contrôlé thermiquement pour turbine à gaz
US5630702A (en) * 1994-11-26 1997-05-20 Asea Brown Boveri Ag Arrangement for influencing the radial clearance of the blading in axial-flow compressors including hollow spaces filled with insulating material
WO1998055738A1 (fr) * 1997-06-05 1998-12-10 Dynatrend Asa Procede relatif au demarrage d'une turbine de puissance et arrangements relatifs a ladite turbine permettant d'eviter les dommages de demarrage subis par le carter et la roue de turbine
US20040219009A1 (en) * 2003-03-06 2004-11-04 Snecma Moteurs Turbomachine with cooled ring segments

Cited By (14)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE102008005482A1 (de) * 2008-01-23 2009-07-30 Rolls-Royce Deutschland Ltd & Co Kg Gasturbine mit einem Verdichter mit selbstheilender Einlaufschicht
US8257016B2 (en) 2008-01-23 2012-09-04 Rolls-Royce Deutschland Ltd & Co Kg Gas turbine with a compressor with self-healing abradable coating
EP2273073A3 (fr) * 2009-06-12 2013-07-03 Rolls-Royce plc Système et procédé pour régler le jeu rotor-stator
US8555477B2 (en) 2009-06-12 2013-10-15 Rolls-Royce Plc System and method for adjusting rotor-stator clearance
EP2351912A1 (fr) * 2010-01-12 2011-08-03 Siemens Aktiengesellschaft Turbine avec système chauffant, centrale énergétique solaire et procédé d'exploitation associés
US8695342B2 (en) 2010-01-12 2014-04-15 Siemens Aktiengesellschaft Heating system for a turbine
FR3002273A1 (fr) * 2013-02-20 2014-08-22 Snecma Dispositif avionique pour la surveillance d'une turbomachine
US9472026B2 (en) 2013-02-20 2016-10-18 Snecma Avionics method and device for monitoring a turbomachine at startup
EP2955327A1 (fr) * 2014-06-10 2015-12-16 Rolls-Royce plc Un ensemble d'étanchéité de pointe d'aube
US9803495B2 (en) 2014-06-10 2017-10-31 Rolls-Royce Plc Assembly
FR3045717A1 (fr) * 2015-12-22 2017-06-23 Snecma Dispositif de pilotage de jeu en sommets d'aubes rotatives de turbine
US10539037B2 (en) 2015-12-22 2020-01-21 Safran Aircraft Engines Device for controlling clearance at the tops of turbine rotating blades
EP3354859A1 (fr) * 2017-01-26 2018-08-01 Safran Aero Boosters SA Système de controle actif de jeu radial de turbomachine et turbomachine associée
BE1024941B1 (fr) * 2017-01-26 2018-08-28 Safran Aero Boosters S.A. Controle actif de jeu pour compresseur de turbomachine

Also Published As

Publication number Publication date
CN1854468A (zh) 2006-11-01
CN1854468B (zh) 2010-11-10
CN101825003A (zh) 2010-09-08
JP2006307853A (ja) 2006-11-09
US20060245910A1 (en) 2006-11-02
EP1717419B1 (fr) 2010-10-13
DE502005010381D1 (de) 2010-11-25
US7766611B2 (en) 2010-08-03
ATE484652T1 (de) 2010-10-15

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