EP1443182A2 - Dispositif de refroidissement - Google Patents

Dispositif de refroidissement Download PDF

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Publication number
EP1443182A2
EP1443182A2 EP04100222A EP04100222A EP1443182A2 EP 1443182 A2 EP1443182 A2 EP 1443182A2 EP 04100222 A EP04100222 A EP 04100222A EP 04100222 A EP04100222 A EP 04100222A EP 1443182 A2 EP1443182 A2 EP 1443182A2
Authority
EP
European Patent Office
Prior art keywords
cavity
component
cross
cooling device
section
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Withdrawn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
EP04100222A
Other languages
German (de)
English (en)
Other versions
EP1443182A3 (fr
Inventor
Christoph Nagler
André Schwind
Ralf Walz
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
Alstom Technology 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 Alstom Technology AG filed Critical Alstom Technology AG
Publication of EP1443182A2 publication Critical patent/EP1443182A2/fr
Publication of EP1443182A3 publication Critical patent/EP1443182A3/fr
Withdrawn legal-status Critical Current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
    • F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
    • F01D11/005—Sealing means between non relatively rotating elements
    • 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
    • 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
    • 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
    • F05D2240/00—Components
    • F05D2240/10—Stators
    • F05D2240/11—Shroud seal segments
    • 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
    • F05D2240/00—Components
    • F05D2240/55—Seals
    • F05D2240/57—Leaf seals
    • 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
    • F05D2260/20—Heat transfer, e.g. cooling
    • F05D2260/201—Heat transfer, e.g. cooling by impingement of a fluid
    • 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
    • F05D2260/60—Fluid transfer
    • F05D2260/607—Preventing clogging or obstruction of flow paths by dirt, dust, or foreign particles

Definitions

  • the present invention relates to a cooling device for loading a first cavity with a cooling gas, in particular in a gas turbine Power plant.
  • a component on a first side is exposed to high thermal stress on its other side cool.
  • a heat shield on the one side is exposed to the hot combustion gases and on his other side exposed to a flow of cooling gas.
  • the respective component have a wall that closes, for example Fastening purposes and the first on this cooled side Separates cavity from a second cavity.
  • the first one can usually be connected to a cooling gas supply Cavity via one or more cooling gas channels from the second cavity be supplied with cooling gas.
  • the part separates the first cavity from a third cavity.
  • forms the third cavity is the hot gas area of a gas turbine.
  • this second component can be another heat shield, a Turbine blade or a seal.
  • the second component can in front of the mouth of the cooling gas duct, which on the one hand Cooling gas mass flow in the first cavity is reduced, so that there can set an undesirable temperature increase. On the other hand, it can come to an undesirable pressure drop in the first cavity, causing the third cavity hot gases bypassing the second component in the first cavity can enter, which also leads to an undesirable Temperature increase in the second cavity leads.
  • the problem described can occur in particular in the case of a gas turbine, if the second component is a seal that can be closed using Holding bolt is held in its target position. Vibrations can do this during operation cause the seal to gnaw into the bolts. In extreme cases they can This weakens the bolts and finally breaks them off. Then not more gasket can slide in front of the cooling gas channel or channels. This results in a deterioration in the cooling effect and a drop in pressure in the first cavity, resulting in an extremely high in a short time Temperature rise in the first cavity can result.
  • the invention seeks to remedy this.
  • the invention as set out in the claims is concerned with the problem for a cooling device of the type mentioned to provide an improved embodiment, the especially when the relative position between the first component changes and the second component an adequate cooling gas supply of the first Cavity.
  • the invention is based on the general idea, the first Cavity-facing mouth region of the cooling gas channel in terms of its Dimensioning and / or positioning to a predetermined Adjust adjustment range within which the relative adjustments move between the two components as expected.
  • This Construction can be used for any possible relative position between the two components A sufficiently large cross-section of the mouth can be made available, so that always an adequate supply of cooling gas to the first cavity and a sufficiently large pressure is available in the first cavity.
  • the performance of the cooling device can be improved.
  • the cooling gas channel can be outside its mouth area have a specified nominal cross section that is smaller than all cross sections in the Mouth area.
  • This nominal cross section forms within the cooling gas duct the narrowest or smallest cross section. Accordingly, by the Nominal cross section in a nominal operating point of the cooling device Cooling gas mass flow through the cooling gas channel and the pressures in the first and defined in the second cavity.
  • Cooling gas mass flow through the cooling gas channel and the pressures in the first and defined in the second cavity can be a minimum cross section with which the muzzle cross section in all provided relative positions of the components is securely open, the same size as or be larger than this nominal cross section. Accordingly, this ensures Construction that in all predicted relative positions between the components the cooling gas mass flow through the cooling gas channel and / or the pressure in the first and the second cavity have the values provided for the nominal operation.
  • the mouth area can have any geometric shape have, which leads to a mouth cross-section that is larger than that Nominal cross-section. Geometries that are easy to produce are preferred.
  • the mouth area can be formed by a chamfer, the the end of the cooling gas channel facing the first cavity is attached.
  • a groove in which several cooling gas channels are provided are in the wall on a system side facing the first cavity a groove can be formed, the at least two cooling gas channels with each other connects, such that the mouth areas of these cooling gas channels through the Groove are formed or pass into this groove.
  • the 1 contains a gas turbine 1, only partially shown, in particular a power plant, a rotor 2, which is not one here shown, rotatable parallel to the cutting plane rotor axis mounted is.
  • the rotor 2 carries blades 3, of which only one in FIG. 1 is shown as an example.
  • the rotor 2 rotates in a housing 4, which several Guide vanes 5 carries, of which only two are shown here.
  • the housing 4 carries a heat shield 6 between two rows of guide vanes, one of which Blade 3 is radially adjacent.
  • the heat shield 6 has a radially inner one with respect to the rotor axis of the rotor 2 inside 7 and radially outside outside 8 on.
  • the first cavity 9 and the second cavity 10 separated from each other by a wall 11, which on the outside 8 of the Heat shield 6 is formed on the heat shield 6 and in the circumferential direction extends.
  • the heat shield 6 On its inside 7, the heat shield 6 is a third cavity 12 exposed, in which the blades 3,5 are arranged, and which in the operation of the Gas turbine 1 is flowed through with hot flow gases. Axially between the Heat shield 6 and a blade root 13 of the downstream adjacent guide blade 5 a gap 14 is formed, through which the first cavity 9 with the third Cavity 12 is connected. Around this connection or around this gap 14 seal, is on a system side facing away from the second cavity 10 15 Wall 11 has a seal 16 arranged on the one hand on the contact side 15 the wall 11 and on the other hand axially supported on the blade root 13. The seal 16 thus separates the first cavity 9 from the third cavity 12.
  • the seal 16 has a U-shaped cross section as an example. It is clear that basically any other cross-sections can be used can, such as a W-shaped cross section or a full cross section or a disc-shaped cross section.
  • Cooling device 17 is provided according to the invention.
  • the second cavity 10 via a cooling gas supply 18 with a cooling gas applied.
  • At least one cooling gas channel 19 is formed in the wall 11, communicating the first cavity 9 with the second cavity 10 combines.
  • the wall 11 usually contains distributed in the circumferential direction several such cooling gas channels 19. Via the cooling gas channels 19 can the cooling gas from the second cavity 10 enter the first cavity 9 and cool the surfaces or components adjacent to the first cavity 9.
  • the first cavity 9 With cooling gas provided.
  • a predetermined pressure is in the first cavity 9 formed, which is advantageously higher than the pressure in the third cavity 12. This ensures that in the event of leaks, no hot gas from the third Cavity 12 enters the first cavity 9.
  • the seal 16 is in the correct operation of the gas turbine 1 approximately in the position shown in Fig. 1, in which it the gas flow through the Cooling gas channel 19 is not hindered. In certain operating situations and / or in the case of (minor) damage, it can happen that the Seal 16 in the radial direction along the wall 11 within a predetermined adjustment range moves. Here, the seal 16 can push one or more cooling channels 19. So that through this Adjustment movement of the seal 16 no impairment of the cooling effect results, is the cooling device 17 with the features of the invention equipped, which is described in more detail below with reference to FIGS. 2 to 9.
  • This mouth region 20 is now inside the wall 11 on the plant side 15 dimensioned and / or positioned so that its mouth cross section 21 protrudes the aforementioned adjustment range of the seal 16, namely to the extent that the mouth cross section 21 in any position Seal 16 is not completely within the adjustment range of the seal 16 can be covered, but always at least with a predetermined
  • the minimum cross section remains open. This minimum cross section is chosen so that A sufficient flow through the cooling gas channel 19 can be guaranteed can, so that on the one hand a sufficient mass flow and on the other hand a sufficient pressure can be provided in the first cavity 9.
  • the seal 16 takes a first extreme position in its Adjustment range in which there is minimal overlap with the Mouth cross section 21 results.
  • This relative position is normal Operating conditions of the gas turbine 1 before. 4, 5 and 8, 9 show a second extreme position of the seal 16 within the adjustment range maximum coverage of the mouth cross section 21. This relative position results in special operating conditions or in the case of calculated damage, for example if a mounting of the seal 16 fails.
  • the default Adjustment range of the seal 16 is in FIGS. 4 and 8 by a double arrow symbolized and labeled 22.
  • the cooling gas channel 19 has a constant Cross section, which is also referred to below as the nominal cross section 23.
  • This nominal cross section 23 is smaller than all cross sections in the Mouth area 20.
  • the nominal cross section 23 defines the nominal operating point of the Gas turbine 1, the cooling gas mass flow through the cooling gas channel 19 and the achievable pressure in the first cavity 9. Furthermore, the Dimensioning of the nominal cross section 23 of the pressure in the second cavity 10 certainly. It is therefore essential for the proper operation of the cooling device 17 not useful, the entire cooling gas channel 19 with the comparatively to equip a large mouth cross section 21. For example, that would be Pressure drop in the second cavity 10 too large.
  • the minimum cross-section is according to an appropriate dimensioning of the mouth cross section 21, which at maximum coverage of the seal 16 remains open, chosen so large that it is at least the same size as the Nominal cross section 23. Accordingly, even with an extreme Adjustment of the seal 16 provided for the nominal operating point Mass flow and the associated pressure conditions in the first cavity 9 and are maintained in the second cavity 10.
  • the cooling gas channel 19 widens in Mouth area 20 to the first cavity 9 until it is Mouth cross section 21 reached.
  • the cooling gas channel 19 tapers from the mouth cross section 21 to Nominal cross section 23. This is achieved, for example, by means of a retrospective attached bevel.
  • the mouth area has 20 also from this cross-sectional expansion 24 to the mouth cross-section 21 a constant cross section.
  • the mouth region 20 be produced by means of a groove 25 which on the contact side 15 into the wall 11 is introduced in such a way that the cooling gas channel 19 in the groove base of the groove 25 opens.
  • the side of the groove 25 that is open toward the first cavity 9 then forms the mouth cross section 21 of the cooling gas channel 19, which is determined by the length of the groove 25 can be designed many times larger than the nominal cross section 23 can.
  • the wall 11 contains a plurality of cooling gas channels 19, it is appropriate to Lay groove 25 so that it spans several, especially all, Cooling gas channels 19 leads.
  • the connected to each other via the groove 25 Cooling gas channels 19 thus have a common, relatively large volume Mouth area 20.
  • the heat shield 6 here forms a first component 6 on which the wall 11 is connected Separation of the first cavity 9 from the second cavity 10 is formed.
  • the seal 16 which forms a second component 16, which on the wall 11th separates the first cavity 9 from the third cavity 12.
  • the second component 16 can also be formed by another component.
  • the blade root 13 can be located directly on the contact side 15 of the wall 11 come to rest and thereby form the second component. It is clear that the present invention is not limited to a heat shield 6, but instead basically on any other component with the corresponding Cooling device 17 is applicable.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP04100222A 2003-01-29 2004-01-22 Dispositif de refroidissement Withdrawn EP1443182A3 (fr)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
DE10303340 2003-01-29
DE10303340A DE10303340A1 (de) 2003-01-29 2003-01-29 Kühleinrichtung

Publications (2)

Publication Number Publication Date
EP1443182A2 true EP1443182A2 (fr) 2004-08-04
EP1443182A3 EP1443182A3 (fr) 2006-12-20

Family

ID=32603010

Family Applications (1)

Application Number Title Priority Date Filing Date
EP04100222A Withdrawn EP1443182A3 (fr) 2003-01-29 2004-01-22 Dispositif de refroidissement

Country Status (3)

Country Link
US (1) US7131814B2 (fr)
EP (1) EP1443182A3 (fr)
DE (1) DE10303340A1 (fr)

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2009103636A1 (fr) * 2008-02-20 2009-08-27 Alstom Technology Ltd. Machine thermique
WO2014189873A2 (fr) * 2013-05-21 2014-11-27 Siemens Energy, Inc. Appareil de refroidissement de segment d'anneau de turbine à gaz
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

Families Citing this family (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US8186933B2 (en) * 2009-03-24 2012-05-29 General Electric Company Systems, methods, and apparatus for passive purge flow control in a turbine
US8371814B2 (en) * 2009-06-24 2013-02-12 Honeywell International Inc. Turbine engine components
US8529193B2 (en) * 2009-11-25 2013-09-10 Honeywell International Inc. Gas turbine engine components with improved film cooling
US8628293B2 (en) 2010-06-17 2014-01-14 Honeywell International Inc. Gas turbine engine components with cooling hole trenches
US20130028704A1 (en) * 2011-07-26 2013-01-31 Thibodeau Anne-Marie B Blade outer air seal with passage joined cavities
US9650900B2 (en) 2012-05-07 2017-05-16 Honeywell International Inc. Gas turbine engine components with film cooling holes having cylindrical to multi-lobe configurations
US10113433B2 (en) 2012-10-04 2018-10-30 Honeywell International Inc. Gas turbine engine components with lateral and forward sweep film cooling holes
US11021965B2 (en) 2016-05-19 2021-06-01 Honeywell International Inc. Engine components with cooling holes having tailored metering and diffuser portions
US10428689B2 (en) * 2017-05-17 2019-10-01 Rolls-Royce Deutschland Ltd & Co Kg Heat shield for a gas turbine engine

Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE1287857B (fr) * 1964-09-24 1969-01-23
DE2918996A1 (de) * 1978-08-14 1980-02-28 Gen Electric Verfahren und vorrichtung zur halterung von dichtungen
US5165847A (en) * 1991-05-20 1992-11-24 General Electric Company Tapered enlargement metering inlet channel for a shroud cooling assembly of gas turbine engines

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FR2548733B1 (fr) * 1983-07-07 1987-07-10 Snecma Dispositif d'etancheite d'aubages mobiles de turbomachine
US5584651A (en) * 1994-10-31 1996-12-17 General Electric Company Cooled shroud
US5641267A (en) * 1995-06-06 1997-06-24 General Electric Company Controlled leakage shroud panel
US5738490A (en) * 1996-05-20 1998-04-14 Pratt & Whitney Canada, Inc. Gas turbine engine shroud seals
US5971703A (en) * 1997-12-05 1999-10-26 Pratt & Whitney Canada Inc. Seal assembly for a gas turbine engine
DE19756734A1 (de) * 1997-12-19 1999-06-24 Bmw Rolls Royce Gmbh Passives Spalthaltungssystem einer Gasturbine
DE19963371A1 (de) * 1999-12-28 2001-07-12 Alstom Power Schweiz Ag Baden Gekühltes Hitzeschild
US6340285B1 (en) * 2000-06-08 2002-01-22 General Electric Company End rail cooling for combined high and low pressure turbine shroud
FR2832178B1 (fr) * 2001-11-15 2004-07-09 Snecma Moteurs Dispositif de refroidissement pour anneaux de turbine a gaz

Patent Citations (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE1287857B (fr) * 1964-09-24 1969-01-23
DE2918996A1 (de) * 1978-08-14 1980-02-28 Gen Electric Verfahren und vorrichtung zur halterung von dichtungen
US5165847A (en) * 1991-05-20 1992-11-24 General Electric Company Tapered enlargement metering inlet channel for a shroud cooling assembly of gas turbine engines

Cited By (7)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2009103636A1 (fr) * 2008-02-20 2009-08-27 Alstom Technology Ltd. Machine thermique
US8272220B2 (en) 2008-02-20 2012-09-25 Alstom Technology Ltd Impingement cooling plate for a hot gas duct of a thermal machine
AU2009216835B2 (en) * 2008-02-20 2013-12-05 General Electric Technology Gmbh Thermal machine
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
WO2014189873A2 (fr) * 2013-05-21 2014-11-27 Siemens Energy, Inc. Appareil de refroidissement de segment d'anneau de turbine à gaz
WO2014189873A3 (fr) * 2013-05-21 2015-01-15 Siemens Energy, Inc. Appareil de refroidissement de segment d'anneau de turbine à gaz
US10233776B2 (en) 2013-05-21 2019-03-19 Siemens Energy, Inc. Gas turbine ring segment cooling apparatus

Also Published As

Publication number Publication date
EP1443182A3 (fr) 2006-12-20
DE10303340A1 (de) 2004-08-26
US20050089396A1 (en) 2005-04-28
US7131814B2 (en) 2006-11-07

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