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EP2031302A1 - Turbine à gaz comprenant un composant refroidissable - Google Patents

Turbine à gaz comprenant un composant refroidissable Download PDF

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Publication number
EP2031302A1
EP2031302A1 EP07016761A EP07016761A EP2031302A1 EP 2031302 A1 EP2031302 A1 EP 2031302A1 EP 07016761 A EP07016761 A EP 07016761A EP 07016761 A EP07016761 A EP 07016761A EP 2031302 A1 EP2031302 A1 EP 2031302A1
Authority
EP
European Patent Office
Prior art keywords
wall
turbulence
hot gas
structural elements
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.)
Withdrawn
Application number
EP07016761A
Other languages
German (de)
English (en)
Inventor
Roland Dr. Liebe
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
Siemens Corp
Original Assignee
Siemens AG
Siemens Corp
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, Siemens Corp filed Critical Siemens AG
Priority to EP07016761A priority Critical patent/EP2031302A1/fr
Publication of EP2031302A1 publication Critical patent/EP2031302A1/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Withdrawn legal-status Critical Current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
    • F01D9/00—Stators
    • F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
    • F01D9/023—Transition ducts between combustor cans and first stage of the turbine in gas-turbine engines; their cooling or sealings
    • 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
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23M—CASINGS, LININGS, WALLS OR DOORS SPECIALLY ADAPTED FOR COMBUSTION CHAMBERS, e.g. FIREBRIDGES; DEVICES FOR DEFLECTING AIR, FLAMES OR COMBUSTION PRODUCTS IN COMBUSTION CHAMBERS; SAFETY ARRANGEMENTS SPECIALLY ADAPTED FOR COMBUSTION APPARATUS; DETAILS OF COMBUSTION CHAMBERS, NOT OTHERWISE PROVIDED FOR
    • F23M5/00—Casings; Linings; Walls
    • F23M5/08—Cooling thereof; Tube walls
    • F23M5/085—Cooling thereof; Tube walls using air or other gas as the cooling medium
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/005—Combined with pressure or heat exchangers
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/10—Two-dimensional
    • F05D2250/18—Two-dimensional patterned
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/28—Three-dimensional patterned
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/20—Three-dimensional
    • F05D2250/28—Three-dimensional patterned
    • F05D2250/283—Three-dimensional patterned honeycomb
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/60—Structure; Surface texture
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/70—Shape
    • F05D2250/71—Shape curved
    • F05D2250/712—Shape curved concave
    • 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
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2212—Improvement of heat transfer by creating turbulence
    • 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
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
    • 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
    • F05D2260/221—Improvement of heat transfer
    • F05D2260/2214—Improvement of heat transfer by increasing the heat transfer surface
    • F05D2260/22141—Improvement of heat transfer by increasing the heat transfer surface using fins or ribs
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/03043—Convection cooled combustion chamber walls with means for guiding the cooling air flow
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/03045—Convection cooled combustion chamber walls provided with turbolators or means for creating turbulences to increase cooling

Definitions

  • the invention relates to a gas turbine with a number of each combined into rows of blades, arranged on a turbine shaft blades and a number of guide vanes combined, connected to a turbine housing vanes and with a number of hot gas components, in which a hot gas flow space bordering outer wall is integrated in each case a number of cooling channels for acting on a cooling medium.
  • Gas turbines are used in many areas to drive generators or work machines.
  • the energy content of a fuel is used to generate a rotational movement of a turbine shaft.
  • the fuel is burned in a combustion chamber, compressed air being supplied by an air compressor.
  • the working medium produced in the combustion chamber by the combustion of the fuel, under high pressure and at high temperature, is guided via a turbine unit arranged downstream of the combustion chamber, where it relaxes to perform work.
  • a number of suitable hot gas-conducting components are provided, for example the combustion chamber per se or its outflow region and this downstream transition piece opening into the turbine unit.
  • a number of rotor blades which are usually combined into blade groups or rows of blades, are arranged thereon and drive the turbine shaft via a momentum transfer from the working medium.
  • To guide the working fluid in the turbine unit also usually connected between adjacent blade rows are connected to the turbine housing Leitschaufelhalln.
  • cooling of the affected components in particular of rotor blades and / or guide vanes of the turbine unit or of the hot gas-conducting components, is usually provided.
  • These affected components in particular the hot gas-carrying components such as combustion chamber walls or transition pieces, are therefore usually designed coolable, in particular, an effective and reliable cooling of the flow space of the hot gas directly limiting component walls should be ensured.
  • the affected outer wall of the respective component usually has a cooling channel integrated into the wall material for the application of a coolant such as, for example, air or steam.
  • a reliable cooling system for the respective component wall can be provided with comparatively simple means, wherein also thermally highly stressed zones of the component are suitable acted upon by coolant.
  • the flow cross-sections achievable for the cooling channels are only limited. Too small a flow cross-section can help, among other things, that the flow resistance during the flow of the coolant through the cooling channel is undesirably high, which can have an effect-reducing effect. In order to limit this disadvantage, the flow rate of the coolant could generally be reduced, which, on the other hand, adversely affects the desired cooling effect again.
  • the invention is therefore an object of the invention to provide a gas turbine of the type mentioned above, in which even with regard to the limited space and a low efficiency for low pressure loss sufficient cooling effect on the thermally highly stressed areas of the respective component can be achieved safely.
  • At least one of the cooling channels integrated into the outer wall of the respective component is provided with turbulence-generating structural elements at its inner wall facing away from the hot gas flow space.
  • the invention is based on the consideration that, in order to ensure a sufficient cooling effect on the mentioned components with acceptable pressure loss, the maximization of the parameter heat transfer / (pressure loss) n should always be taken into account as a design target for the respective cooling channels, the exponent n being one of the dependent component and their specific conditions of use is dependent coefficient.
  • the maximization of this characteristic as the design target can be promoted by maximizing the heat transfer coefficient of the surface of the respective cooling channel.
  • the inner wall of the respective cooling channel should be provided with turbulence-generating structures in order to generate a particularly intimate contact between the coolant and the channel wall and thus a particularly effective heat exchange between the channel wall by the turbulence thus flowing in the coolant channel and to ensure coolant.
  • this high heat transfer is achievable, without at the same time an excessively high pressure loss in the coolant is to be accepted.
  • a particularly thin channel wall between the cooling channel and the hot gas flow space can be achieved if that wall of the cooling channel which faces the hot gas flow space is free of turbulence-generating structures.
  • the desired high heat transfer between the coolant and the channel wall can be achieved in an advantageous embodiment by providing mixer fins, rib structures or other suitably selected structural elements projecting from the wall material into the flow region of the coolant as turbulence-generating structural elements to the respective channel wall or wall zone.
  • mixer fins, rib structures or other suitably selected structural elements projecting from the wall material into the flow region of the coolant as turbulence-generating structural elements to the respective channel wall or wall zone.
  • local wall recesses also referred to as so-called “dimples" or concave indentations in the wall surface, are provided as turbulence-generating structural elements.
  • Such local wall recesses or “dimples” may have various outer contours such as circular, square, triangular, hexagonal or the like, various basic shapes such as spherical, cylindrical or the like, and geometry parameters such as embossing depths, effective surface area and the like.
  • impressions or local wall recesses have the particular advantage that Such structures, the specific heat transfer at the interface can be significantly increased, without thereby increasing the pressure loss in particular measure.
  • the pressure loss to be taken into account for a desired increase in the heat transfer through such local wall depressions is usually only about half as large as in the case of structures protruding from the wall surface.
  • a particularly uniform flow course and thus a particularly uniform heat transfer with limited pressure loss, also with regard to the swirling or turbulence generated as intended in the flowing coolant, can be achieved by arranging the turbulence-generating structural elements in a further advantageous embodiment in an associated surface region of the inner wall in a regular grid arrangement.
  • a grid arrangement a hexagonal grid is provided in a further advantageous embodiment, in which a particularly high packing or area density of the structural elements can be achieved.
  • the turbulence-generating structural elements are adapted in a further advantageous embodiment in its outer contour to the grid structure. When selecting a hexagonal grid for the grid arrangement, a hexagonal outer contour for the structural elements is thus provided in this embodiment.
  • said turbulence-generating structuring means are used in a cooling channel of a combustion chamber wall or of a transition piece of the gas turbine connected downstream of the combustion gas side on the hot gas side.
  • the advantages achieved by the invention are in particular that by attaching turbulence-generating structural elements and in particular of concave wall depressions on the inner wall of an integrated into the outer wall of the hot gas leading component cooling channel with limited pressure loss, a particularly high heat transfer and thus even with limited space for the cooling system a particularly effective cooling of thermally highly loaded zones can be guaranteed.
  • by the arrangement of these structural elements on the side facing away from the hot gas flow chamber inner wall of the cooling channel is also an excessively high temperature gradient in the wall material immediately adjacent to the flow space of the hot gas safely avoided.
  • the gas turbine 1 has a compressor 2 for combustion air, a combustion chamber 4 and a turbine 6 for driving the compressor 2 and a generator, not shown or a work machine.
  • the turbine 6 and the compressor 2 are arranged on a common, also called turbine rotor turbine shaft 8, with which the generator or the working machine is connected, and which is rotatably mounted about its central axis 9.
  • the combustion chamber 4 is equipped with a number of burners 10 for the combustion of a liquid or gaseous fuel.
  • the turbine 6 has a number of rotatable blades 12 connected to the turbine shaft 8.
  • the blades 12 are arranged in a ring on the turbine shaft 8 and thus form a number of blade rows.
  • the turbine 6 comprises a number of fixed vanes 14, which are also fixed in a ring shape with the formation of rows of vanes on an inner casing 16 of the turbine 6.
  • the blades 12 serve to drive the turbine shaft 8 by momentum transfer from the turbine 6 flowing through the working medium M.
  • the vanes 14, however, serve to guide the flow of the working medium M between two seen in the flow direction of the working medium M consecutive blade rows or blade rings.
  • a successive pair of a ring of vanes 14 or a row of vanes and a ring of blades 12 or a blade row is also referred to as a turbine stage.
  • Each vane 14 has a platform 18, also referred to as a blade root, which is arranged to fix the respective vane 14 on the inner housing 16 of the turbine 6 as a wall element.
  • the platform 18 is a thermally comparatively heavily loaded component, which forms the outer boundary of a hot gas channel for the working medium M flowing through the turbine 6.
  • Each blade 12 is attached to the turbine shaft 8 in an analogous manner via a platform 20, also referred to as a blade root.
  • each guide ring 21 on the inner housing 16 of the turbine 6 is arranged between the spaced-apart platforms 18 of the guide vanes 14 of two adjacent rows of guide vanes.
  • the outer surface of each guide ring 19 is also exposed to the hot, the turbine 6 flowing through the working fluid M and spaced in the radial direction from the outer end 22 of the blade 12 opposite it through a gap.
  • the guide rings 19 arranged between adjacent guide blade rows serve in particular as cover elements which protect the inner wall 16 or other housing mounting parts from thermal overload by the hot working medium M flowing through the turbine 6.
  • the gas turbine 1 is designed for a comparatively high outlet temperature of the working medium M emerging from the combustion chamber 4 from about 1200 ° C. to 1300 ° C.
  • the gas turbine 1 is designed for a comparatively high outlet temperature of the working medium M emerging from the combustion chamber 4 from about 1200 ° C. to 1300 ° C.
  • the respective Cooling channel 32 is provided on its inner wall with turbulence-generating structural elements 34, which increase the heat transfer from guided in the cooling passage 32 coolant into the wall with only a limited increase in the pressure loss.
  • these turbulence-generating structure elements 34 are arranged in the embodiment exclusively on the side facing away from the hot gas flow chamber 36 inner wall 38 of the cooling channel 32.
  • the inner wall 40 of the cooling channel 32 facing the hot gas flow space 36 is kept free of such structural elements, so that an excessively high temperature gradient in the region between the hot gas flow space 36 and the inner wall 40 of the cooling channel 32 facing the latter can be avoided.
  • the inner wall 40 of the cooling channel 32 is smooth or even.
  • the turbulence-generating structure elements 34 are configured in the embodiment as local wall recesses or concave depressions 42, so-called “dimples".
  • these local depressions can have any suitable outer contours such as, for example, circular, square or the like.
  • the recesses are provided with a hexagonal or hexagonal outer contour.
  • the recesses 42 are arranged in a regular grid structure on the surface, wherein in the embodiment, a hexagonal grid structure is selected.
  • the recesses 42 are adapted in their outer contour to the grid structure.

Landscapes

  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP07016761A 2007-08-27 2007-08-27 Turbine à gaz comprenant un composant refroidissable Withdrawn EP2031302A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP07016761A EP2031302A1 (fr) 2007-08-27 2007-08-27 Turbine à gaz comprenant un composant refroidissable

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP07016761A EP2031302A1 (fr) 2007-08-27 2007-08-27 Turbine à gaz comprenant un composant refroidissable

Publications (1)

Publication Number Publication Date
EP2031302A1 true EP2031302A1 (fr) 2009-03-04

Family

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EP07016761A Withdrawn EP2031302A1 (fr) 2007-08-27 2007-08-27 Turbine à gaz comprenant un composant refroidissable

Country Status (1)

Country Link
EP (1) EP2031302A1 (fr)

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2015089278A1 (fr) * 2013-12-12 2015-06-18 Siemens Energy, Inc. Section de combustion pour un moteur à turbine à gaz
EP3067622A1 (fr) 2015-03-12 2016-09-14 General Electric Technology GmbH Chambre de combustion à double paroi

Citations (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2648519A (en) * 1948-04-22 1953-08-11 Campini Secondo Cooling combustion turbines
US4064300A (en) * 1975-07-16 1977-12-20 Rolls-Royce Limited Laminated materials
US4184326A (en) * 1975-12-05 1980-01-22 United Technologies Corporation Louver construction for liner of gas turbine engine combustor
US4642993A (en) * 1985-04-29 1987-02-17 Avco Corporation Combustor liner wall
EP0974735A2 (fr) * 1998-07-20 2000-01-26 General Electric Company Chicane avec des protubérances
US6098397A (en) * 1998-06-08 2000-08-08 Caterpillar Inc. Combustor for a low-emissions gas turbine engine
EP1143107A2 (fr) * 2000-04-06 2001-10-10 General Electric Company Refroidissement de la bride d'un canal de transition
US6408628B1 (en) * 1999-11-06 2002-06-25 Rolls-Royce Plc Wall elements for gas turbine engine combustors
EP1394389A1 (fr) * 2002-08-30 2004-03-03 General Electric Company Echangeur de chaleur amélioré d'un équipement de génération d'énergie
EP1400750A2 (fr) * 2002-09-18 2004-03-24 General Electric Company Chemise de chambre de combustion à double paroi avec des canaux de refroidissement
DE10248548A1 (de) * 2002-10-18 2004-04-29 Alstom (Switzerland) Ltd. Kühlbares Bauteil
EP1628076A1 (fr) * 2004-08-13 2006-02-22 Siemens Aktiengesellschaft Canal de refroidissement, chambre de combustion et turbine à gaz
DE102005038395A1 (de) * 2004-08-26 2006-03-02 General Electric Co. Brennkammerkühlung mit geneigten segmentierten Flächen

Patent Citations (13)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2648519A (en) * 1948-04-22 1953-08-11 Campini Secondo Cooling combustion turbines
US4064300A (en) * 1975-07-16 1977-12-20 Rolls-Royce Limited Laminated materials
US4184326A (en) * 1975-12-05 1980-01-22 United Technologies Corporation Louver construction for liner of gas turbine engine combustor
US4642993A (en) * 1985-04-29 1987-02-17 Avco Corporation Combustor liner wall
US6098397A (en) * 1998-06-08 2000-08-08 Caterpillar Inc. Combustor for a low-emissions gas turbine engine
EP0974735A2 (fr) * 1998-07-20 2000-01-26 General Electric Company Chicane avec des protubérances
US6408628B1 (en) * 1999-11-06 2002-06-25 Rolls-Royce Plc Wall elements for gas turbine engine combustors
EP1143107A2 (fr) * 2000-04-06 2001-10-10 General Electric Company Refroidissement de la bride d'un canal de transition
EP1394389A1 (fr) * 2002-08-30 2004-03-03 General Electric Company Echangeur de chaleur amélioré d'un équipement de génération d'énergie
EP1400750A2 (fr) * 2002-09-18 2004-03-24 General Electric Company Chemise de chambre de combustion à double paroi avec des canaux de refroidissement
DE10248548A1 (de) * 2002-10-18 2004-04-29 Alstom (Switzerland) Ltd. Kühlbares Bauteil
EP1628076A1 (fr) * 2004-08-13 2006-02-22 Siemens Aktiengesellschaft Canal de refroidissement, chambre de combustion et turbine à gaz
DE102005038395A1 (de) * 2004-08-26 2006-03-02 General Electric Co. Brennkammerkühlung mit geneigten segmentierten Flächen

Cited By (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2015089278A1 (fr) * 2013-12-12 2015-06-18 Siemens Energy, Inc. Section de combustion pour un moteur à turbine à gaz
EP3067622A1 (fr) 2015-03-12 2016-09-14 General Electric Technology GmbH Chambre de combustion à double paroi

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