EP1048822A2 - Bouclier thermique pour turbine à gaz - Google Patents

Bouclier thermique pour turbine à gaz Download PDF

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Publication number
EP1048822A2
EP1048822A2 EP00810216A EP00810216A EP1048822A2 EP 1048822 A2 EP1048822 A2 EP 1048822A2 EP 00810216 A EP00810216 A EP 00810216A EP 00810216 A EP00810216 A EP 00810216A EP 1048822 A2 EP1048822 A2 EP 1048822A2
Authority
EP
European Patent Office
Prior art keywords
heat shield
cooling
segments
cooling air
gas turbine
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
EP00810216A
Other languages
German (de)
English (en)
Other versions
EP1048822B1 (fr
EP1048822A3 (fr
Inventor
Christoph Nagler
Christof Pfeiffer
Ulrich Wellenkamp
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.)
General Electric Technology GmbH
Original Assignee
ABB Alstom Power Switzerland Ltd
Alstom Schweiz 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 ABB Alstom Power Switzerland Ltd, Alstom Schweiz AG filed Critical ABB Alstom Power Switzerland Ltd
Publication of EP1048822A2 publication Critical patent/EP1048822A2/fr
Publication of EP1048822A3 publication Critical patent/EP1048822A3/fr
Application granted granted Critical
Publication of EP1048822B1 publication Critical patent/EP1048822B1/fr
Anticipated expiration legal-status Critical
Expired - Lifetime 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
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/08—Cooling; Heating; Heat-insulation
    • F01D25/12—Cooling
    • 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

Definitions

  • the present invention relates to the field of gas turbine technology. It relates to a heat shield for a gas turbine, which heat shield is used in the Hot gas duct of the gas turbine rotating blades of a stage of the gas turbine encloses annularly and from a plurality of one behind the other in the circumferential direction arranged, curved in a segment of a circle and cooled from the outside
  • Such a heat shield is e.g. from the publications US-A-4,177,004, US-A-4,551,064, US-A-5,071,313, US-A-5,584,651 or EP-A1-0 516 322 are known.
  • Heat shields for gas turbines which form the rotor blades of a turbine stage surround and on the one hand limit the hot gas duct to the outside and on the other hand, the gap between the outer wall of the hot gas duct and the ends keep the blades as small as possible for reasons of efficiency, without causing a sliding contact at changing temperatures, have been known for a long time.
  • Such heat shields usually consist of one Variety of heat shield segments curved in the shape of a segment of a circle, extending in the circumferential direction arranged one behind the other form a closed ring.
  • the individual heat shield segments are often releasably attached to a carrier, that concentrically surrounds the heat shield.
  • a carrier that concentrically surrounds the heat shield.
  • the heat shield or the individual heat shield segments are in operation exposed to high thermal stress on the gas turbine.
  • This thermal Load can have a negative impact on the heat shield itself to have.
  • the heat can escape through the shield and cause damage there. Therefore, precautions are usually taken hit the heat shield segments from the back or outside forth by compressed cooling air, which mostly comes from the compressor part of the Gas turbine or the plenum comes to cool in a suitable manner.
  • This cooling should be as uniform and efficient as possible and all polluted areas of the heat shield.
  • it should be prevented Hot gas in the adjacent column in the outer wall of the hot gas duct penetrates and the underlying parts of the construction in undesirable Way heated.
  • US-A-4,177,004 discloses a heat shield for a gas turbine (there Fig. 1, 2 and 4), in which only on the downstream longitudinal side of the heat shield segments Cooling air from the cavity (52) behind it through cooling holes (66) is sent into the adjacent space (48) and from there is passed through cooling grooves (67) in the bracket part (43) into the hot gas duct (Fig. 4, Fig. 5).
  • the upstream long side of the heat shield segment (Fig. 3) is only externally washed by cooling air, which is in other ways flows into the cavity (62) behind it.
  • This arrangement has the Disadvantage that the overall heat shield segment is cooled unevenly, because on the upstream side of the heat shield segment Cooling from the rear practically does not take place.
  • Another disadvantage is that the cooling grooves (67) have been introduced into the clamp element (43), which leads to a considerable additional effort in terms of manufacturing technology.
  • a heat shield is disclosed in US Pat. No. 5,584,651, the segments of which are shown in FIG an inner cavity (38) is formed upstream edge (FIG. 2), flows through the cooling air and through outlet bores arranged directly on the edge (44) emerges into the hot gas duct.
  • the downstream ones are particularly affected inner arms of the heat shield segments with the edges (28b in Fig. 1).
  • the object is achieved by the entirety of the features of claim 1.
  • the essence of the invention is, on both long sides of the heat shields, So both upstream and downstream, from behind the segments lying cavity cooling air through appropriate cooling holes in the adjacent Lead column and so the two at the same time and evenly Longitudinal areas of the heat shield segments to cool and the gaps against flooding the penetration of hot gases. All cooling and flooding devices are there (in the form of cooling holes or cooling grooves) on the heat shield segment self-arranged, which makes production much easier and an adaptation of the remaining parts of the hot gas duct is unnecessary.
  • the drain the cooling air on both long sides of the heat shield segments also has Consequence that the cooling air is more uniform over the delimiting the cavity Outside of the segments strokes and so the entire segment area evenly cools. As a result, the thermal load over the entire surface evenly reduced and the service life of the heat shield segments significantly extended.
  • a first preferred embodiment of the heat shield according to the invention is characterized in that the heat shield segments by means of brackets on Brackets are attached, which brackets are bent inward with an L-shape Ends from both sides under the beam into those formed between the pairs of arms Intervening spaces intervene that the flowing out of the cooling holes Cooling air in the spaces between the L-bends bent inwards Ends of the brackets and the inner arms of the heat shield segments the columns is guided, and that to guide those emerging from the cooling holes Cooling air in the outside of the inner arms to the cooling holes aligned cooling grooves are embedded.
  • Through the cooling grooves in the inner Arms increase the heat transfer surface on the arms and cooling the arms (farthest from the cooling air-filled cavity) are essential equalized and improved.
  • a second preferred embodiment of the heat shield according to the invention is characterized in that to reduce the deflection of the heat shield with temperature changes on the outside of the heat shield segments in Area of the cavity arranged axially extending stiffening ribs or are molded on that within the cavity and from the outside of the heat shield segments spaced a circumferential, with openings provided baffle plate is arranged, and that within the stiffening ribs individual noses or pins protruding radially outwards are on which the baffle plate rests.
  • the stiffening ribs with the molded noses stiffen the heat shield segments in the axial direction and this reduces the risk of the blades rubbing against the heat shield. They also improve the heat transfer between the segment and the cooling air flowing through the cavity.
  • the noses used to support the Baffle cooling plates can serve together with the stiffening ribs simple way to be molded when casting the segments.
  • FIG. 1 the partially longitudinally sectioned arrangement of a Heat shield in a gas turbine 10 according to a first preferred embodiment presented the invention.
  • the figure shows a section of the (rotationally symmetrical) hot gas duct 11 of the gas turbine, which of the hot ones Combustion gases from the combustion chamber (not shown) of the gas turbine is flowed through in the direction of the four parallel arrows.
  • the Hot gas duct 11 are arranged guide vanes 13, which are in the radial direction extend and at its outer end into an outer ring 14, the the hot gas channel 11 in the area of the guide vanes 13 is limited to the outside.
  • rotor blades 12 Downstream of the guide vanes 13 are rotor blades 12 which rest on a (Not shown) rotor of the gas turbine are attached and together with this rotate about the turbine axis when they are flowing with the one in the hot gas duct 11 Hot gas are applied. Behind the ring of blades 12 can follow further guide vane and rotor blade rings downstream, on the no further reference needs to be made here. In any case, the hot gas duct 11 behind the blades 12 outwards through an intermediate ring 15 or limited by a subsequent guide vane.
  • the ring of the blades 12 is surrounded concentrically by a heat shield, which is made up of a large number of individual curved segments, in Heat shield segments 17 arranged one behind the other in the circumferential direction.
  • a heat shield segment 17 is within the overall arrangement in FIG. 1 and reproduced in cross section in FIG. The Overall, the heat shield delimits the hot gas duct 11 in the region of the rotor blades 12 and simultaneously determines the gap between the channel wall and the outer end of the blades 12.
  • the individual heat shield segments 17 are curved plates attached to their Long sides, i.e. those oriented transversely to the flow direction or to the turbine axis Pages running circumferentially, possibly with incisions provided rails, each having a pair of axially projecting, parallel and spaced arms 21, 22 and 23, 24 (see also the comparable Fig. 3 of US-A-5,071,313).
  • the heat shield segments 17 are to form a cavity 20 on the inside a concentrically rotating, annular carrier 16 attached.
  • the Attachment takes place via two brackets 18 and 19, which are L-shaped inside bent ends from both sides under the carrier 16 in the between the arm pairs 21, 22 and 23, 24 formed spaces 25 and 26 respectively intervention. To have enough play for different thermal expansion too have between the brackets 18 and 19 and the adjacent Wall elements 15 and 14 radial gaps 29 and 30 left free.
  • the heat shield segments 17 are cooled from the outside via the cavity 20. Compressed cooling air is placed in this cavity at a location (not shown) admitted from the plenum of the gas turbine, which then through on both long sides of the heat shield segment 17 arranged cooling holes 27, 28 in the Spaces 25 and 26 between the pairs of arms 21, 22 and 23, 24 flows out (See the curved arrows in cavity 20 of FIG. 1).
  • the cooling bores 27, 28 are arranged so that the cooling air between the inside (bottom) the L-shaped bent ends of the brackets 18, 19 and the outer sides (Top) of the inner arms 21, 23 out into the gap 29 and 30 flows and exits from there into the hot gas channel 11.
  • Fig. 3 shows these cooling grooves 31, 32 in plan view 4 and 5 show the cooling grooves or cooling holes in cross section.
  • the described type of cooling air routing has several requirements fulfilled safely and easily: Because the cooling air is even on both Longitudinal sides emerges from the cavity 20, the bottom of the cavity 20 or the outside of the heat shield segment evenly over the entire surface Cooling air is applied so that local overheating can be safely avoided. At the same time, too much heat is prevented by conduction into the outside Arms 22, 24 and from there into the carrier. Furthermore, the Brackets 18, 19 effectively cooled at their angled end, so that they too conduct little heat to the outside. In addition, the inside are also Arms 21, 23 effectively protected against overheating. Eventually through the exiting cooling air, the column 29, 30 flooded with cooling air, whereby a undesired penetration of hot gas into the column is reliably avoided.
  • Cooling holes 27, 28 and the cooling grooves 31, 32 aligned therewith - as from the 3 - in the plane of the heat shield segment 17th from the axial direction to the direction of rotation 42 of the rotor blade 12 or Gas turbine are arranged tilted.
  • the position of the heat shield segments 17 determines significantly the gap between the heat shield and the outer end of the blades 12. On the one hand, this gap should be as small as possible, efficiency losses to minimize. On the other hand, the gap must be large enough to allow for different temperatures and the associated different expansions of the elements a sliding contact between the blades and Avoid heat shield if possible. In order to keep the tolerances tight, it is advantageous to bend the heat shield segments due to temperature to reduce that according to FIGS. 6 to 10 on the outside of the Heat shield segments 17 'running from one to the other longitudinal side axial Stiffening ribs 33 are arranged. These stiffening ribs 33 can are advantageously molded on when the heat shield segments 17 'are cast.
  • baffle cooling plate 36 can thus be close to the outside without special shaping the heat shield segments 17 'are placed, whereby the cooling effect of the cooling air flowing through the openings 37 in the impingement cooling plate 36 increases significantly becomes.
  • the lugs or pins 34 increase the heat transfer area and provide additional swirling of the cooling air.
  • a further improvement of the cooling can be achieved or a local one Prevent overheating due to an undesirable cooling air leak if undesirable Cooling air losses can be effectively limited or avoided entirely.
  • axial elastic seals 39, 41 are provided, which drain the Cooling air flowing out of the cooling bores 27, 28 into the gap between the Brackets 18, 19 and the carrier 16 prevented. Because the cooling air on the seals 39 passed directly, the seals are effectively cooled at the same time. Additional axial elastic seals 38, 40, which between the brackets 18, 19th and the carrier 16 are arranged, further improve the seal.
  • this sealed arrangement consists in preventing that hot gas can break in and lead to local overheating.
  • the cooling air leakage is minimized and the cooling air at the places for cooling used where it is actually required. The reduced leakage and The targeted use of cooling air leads to an improvement in efficiency the turbine stage or the machine as a whole.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP00810216A 1999-04-29 2000-03-15 Bouclier thermique pour turbine à gaz Expired - Lifetime EP1048822B1 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE19919654A DE19919654A1 (de) 1999-04-29 1999-04-29 Hitzeschild für eine Gasturbine
DE19919654 1999-04-29

Publications (3)

Publication Number Publication Date
EP1048822A2 true EP1048822A2 (fr) 2000-11-02
EP1048822A3 EP1048822A3 (fr) 2002-07-31
EP1048822B1 EP1048822B1 (fr) 2006-05-17

Family

ID=7906382

Family Applications (1)

Application Number Title Priority Date Filing Date
EP00810216A Expired - Lifetime EP1048822B1 (fr) 1999-04-29 2000-03-15 Bouclier thermique pour turbine à gaz

Country Status (3)

Country Link
US (1) US6302642B1 (fr)
EP (1) EP1048822B1 (fr)
DE (2) DE19919654A1 (fr)

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1225305A2 (fr) * 2001-01-19 2002-07-24 Mitsubishi Heavy Industries, Ltd. Segment de virole pour turbine à gaz
EP1927725A2 (fr) * 2006-11-30 2008-06-04 General Electric Company Système pour faciliter le refroidissement de film récupéré distribué selon la préférence pour un assemblage d'anneaux de turbine
WO2008128876A1 (fr) * 2007-04-19 2008-10-30 Alstom Technology Ltd Écran thermique de stator
CN101737103B (zh) * 2008-11-05 2014-12-17 通用电气公司 关于护罩冷却的方法及设备
FR3112806A1 (fr) * 2020-07-23 2022-01-28 Safran Aircraft Engines Couronne de maintien de secteurs d’étanchéité d’une turbine basse pression

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Publication number Priority date Publication date Assignee Title
DE19938443A1 (de) * 1999-08-13 2001-02-15 Abb Alstom Power Ch Ag Befestigungs- und Fixierungsvorrichtung
DE19945581B4 (de) * 1999-09-23 2014-04-03 Alstom Technology Ltd. Turbomaschine
CA2372984C (fr) * 2000-03-07 2005-05-10 Mitsubishi Heavy Industries, Ltd. Anneau fendu de turbine a gaz
GB2378730B (en) * 2001-08-18 2005-03-16 Rolls Royce Plc Cooled segments surrounding turbine blades
US6783324B2 (en) * 2002-08-15 2004-08-31 General Electric Company Compressor bleed case
US6899518B2 (en) 2002-12-23 2005-05-31 Pratt & Whitney Canada Corp. Turbine shroud segment apparatus for reusing cooling air
JP4191552B2 (ja) * 2003-07-14 2008-12-03 三菱重工業株式会社 ガスタービン尾筒の冷却構造
US7165937B2 (en) * 2004-12-06 2007-01-23 General Electric Company Methods and apparatus for maintaining rotor assembly tip clearances
US7520715B2 (en) * 2005-07-19 2009-04-21 Pratt & Whitney Canada Corp. Turbine shroud segment transpiration cooling with individual cast inlet and outlet cavities
US20070020088A1 (en) * 2005-07-20 2007-01-25 Pratt & Whitney Canada Corp. Turbine shroud segment impingement cooling on vane outer shroud
US7278820B2 (en) * 2005-10-04 2007-10-09 Siemens Power Generation, Inc. Ring seal system with reduced cooling requirements
US7334985B2 (en) * 2005-10-11 2008-02-26 United Technologies Corporation Shroud with aero-effective cooling
US7740442B2 (en) * 2006-11-30 2010-06-22 General Electric Company Methods and system for cooling integral turbine nozzle and shroud assemblies
US7740444B2 (en) * 2006-11-30 2010-06-22 General Electric Company Methods and system for cooling integral turbine shround assemblies
US7665953B2 (en) * 2006-11-30 2010-02-23 General Electric Company Methods and system for recuperated cooling of integral turbine nozzle and shroud assemblies
US8240980B1 (en) 2007-10-19 2012-08-14 Florida Turbine Technologies, Inc. Turbine inter-stage gap cooling and sealing arrangement
US8251637B2 (en) * 2008-05-16 2012-08-28 General Electric Company Systems and methods for modifying modal vibration associated with a turbine
US8556575B2 (en) * 2010-03-26 2013-10-15 United Technologies Corporation Blade outer seal for a gas turbine engine
RU2543101C2 (ru) * 2010-11-29 2015-02-27 Альстом Текнолоджи Лтд Осевая газовая турбина
EP2508713A1 (fr) * 2011-04-04 2012-10-10 Siemens Aktiengesellschaft Turbine à gaz comprenant un écran thermique et procédé d'opération
US9574455B2 (en) * 2012-07-16 2017-02-21 United Technologies Corporation Blade outer air seal with cooling features
EP2949873A1 (fr) * 2014-05-27 2015-12-02 Siemens Aktiengesellschaft Turbomachine avec blindage à l'ingestion et utilisation de la turbomachine
US10400619B2 (en) 2014-06-12 2019-09-03 General Electric Company Shroud hanger assembly
FR3036436B1 (fr) 2015-05-22 2020-01-24 Safran Ceramics Ensemble d'anneau de turbine avec maintien par brides
US10221713B2 (en) * 2015-05-26 2019-03-05 Rolls-Royce Corporation Shroud cartridge having a ceramic matrix composite seal segment
EP3179053B1 (fr) 2015-12-07 2019-04-03 MTU Aero Engines GmbH Structure de carter de turbomachine avec écran de protection thermique
US20170198602A1 (en) * 2016-01-11 2017-07-13 General Electric Company Gas turbine engine with a cooled nozzle segment
KR101965500B1 (ko) * 2017-09-11 2019-04-03 두산중공업 주식회사 터빈의 블레이드 시일 구조 및 이를 포함하는 터빈 및 가스터빈

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US4551064A (en) 1982-03-05 1985-11-05 Rolls-Royce Limited Turbine shroud and turbine shroud assembly
US5071313A (en) 1990-01-16 1991-12-10 General Electric Company Rotor blade shroud segment
EP0516322A1 (fr) 1991-05-20 1992-12-02 General Electric Company Refroidissement pour anneau de stator de turbine à gaz
US5584651A (en) 1994-10-31 1996-12-17 General Electric Company Cooled shroud

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Publication number Priority date Publication date Assignee Title
US4177004A (en) 1977-10-31 1979-12-04 General Electric Company Combined turbine shroud and vane support structure
US4551064A (en) 1982-03-05 1985-11-05 Rolls-Royce Limited Turbine shroud and turbine shroud assembly
US5071313A (en) 1990-01-16 1991-12-10 General Electric Company Rotor blade shroud segment
EP0516322A1 (fr) 1991-05-20 1992-12-02 General Electric Company Refroidissement pour anneau de stator de turbine à gaz
US5584651A (en) 1994-10-31 1996-12-17 General Electric Company Cooled shroud

Cited By (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
EP1225305A2 (fr) * 2001-01-19 2002-07-24 Mitsubishi Heavy Industries, Ltd. Segment de virole pour turbine à gaz
EP1225305A3 (fr) * 2001-01-19 2006-05-17 Mitsubishi Heavy Industries, Ltd. Segment de virole pour turbine à gaz
EP1927725A2 (fr) * 2006-11-30 2008-06-04 General Electric Company Système pour faciliter le refroidissement de film récupéré distribué selon la préférence pour un assemblage d'anneaux de turbine
JP2008138667A (ja) * 2006-11-30 2008-06-19 General Electric Co <Ge> タービンシュラウドアセンブリの優先配分復熱式フィルム冷却を容易にするシステム
EP1927725A3 (fr) * 2006-11-30 2010-03-10 General Electric Company Système pour faciliter le refroidissement de film récupéré distribué selon la préférence pour un assemblage d'anneaux de turbine
WO2008128876A1 (fr) * 2007-04-19 2008-10-30 Alstom Technology Ltd Écran thermique de stator
US7997856B2 (en) 2007-04-19 2011-08-16 Alstom Technology Ltd. Stator heat shield
CN101737103B (zh) * 2008-11-05 2014-12-17 通用电气公司 关于护罩冷却的方法及设备
FR3112806A1 (fr) * 2020-07-23 2022-01-28 Safran Aircraft Engines Couronne de maintien de secteurs d’étanchéité d’une turbine basse pression

Also Published As

Publication number Publication date
US6302642B1 (en) 2001-10-16
DE50012746D1 (de) 2006-06-22
EP1048822B1 (fr) 2006-05-17
EP1048822A3 (fr) 2002-07-31
DE19919654A1 (de) 2000-11-02

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