EP2261565A1 - Chambre de combustion de turbine à gaz et turbine à gaz - Google Patents

Chambre de combustion de turbine à gaz et turbine à gaz Download PDF

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Publication number
EP2261565A1
EP2261565A1 EP09162297A EP09162297A EP2261565A1 EP 2261565 A1 EP2261565 A1 EP 2261565A1 EP 09162297 A EP09162297 A EP 09162297A EP 09162297 A EP09162297 A EP 09162297A EP 2261565 A1 EP2261565 A1 EP 2261565A1
Authority
EP
European Patent Office
Prior art keywords
wall
gas turbine
turbine combustor
combustion chamber
inlet openings
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Withdrawn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
EP09162297A
Other languages
German (de)
English (en)
Inventor
Björn Burbach
Christoph Buse
Alessandro Casu
Giacomo Colmegna
Uwe Gruschka
Birgit Grüger
Andreas Heilos
Thomas Alexis Schneider
Werner Stamm
Stefan Völker
Ulrich Wörz
Adam Zimmermann
Tilman Auf Dem Kampe
Jaap Van Kampen
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Siemens AG
Original Assignee
Siemens AG
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Siemens AG filed Critical Siemens AG
Priority to EP09162297A priority Critical patent/EP2261565A1/fr
Publication of EP2261565A1 publication Critical patent/EP2261565A1/fr
Withdrawn legal-status Critical Current

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Classifications

    • 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
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/44—Combustion chambers comprising a single tubular flame tube within a tubular casing
    • 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
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04—Air inlet arrangements
    • F23R3/06—Arrangement of apertures along the flame tube
    • F23R3/08—Arrangement of apertures along the flame tube between annular flame tube sections, e.g. flame tubes with telescopic sections
    • 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
    • F23M2900/00—Special features of, or arrangements for combustion chambers
    • F23M2900/05004—Special materials for walls or lining
    • 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
    • F23R2900/03042—Film cooled combustion chamber walls or domes

Definitions

  • the present invention relates to a gas turbine combustor having a substantially rotationally symmetrical cross section and at least one axial section, which has an inner wall with an outer side and an outer wall with an inner side facing away from the inner wall and spaced from the inner side, so that between the outer side and the Inside a at least one cooling fluid channel forming space is present.
  • the present invention relates to a gas turbine.
  • a gas turbine comprises as essential components a compressor, a turbine with blades and guide vanes and at least one combustion chamber.
  • the blades of the turbine are arranged on a shaft extending mostly through the entire gas turbine, which is coupled to a consumer, such as a generator for power generation.
  • the shaft provided with the blades is also called turbine runner or rotor.
  • the combustion chamber is supplied with compressed air from the compressor.
  • the compressed air is mixed with a fuel, such as oil or gas, and burned in the combustion chamber.
  • the hot combustion gases are finally fed via a combustion chamber outlet of the turbine, where they transmit momentum to the blades under relaxation and cooling and thus do work.
  • combustion chambers of so-called diffusion combustion systems in which a fuel-rich fuel-air mixture is burned, are exposed to very high temperatures during operation of the gas turbine.
  • the combustion chamber is in this case a mechanical container, which serves to stabilize the flame and to ensure the transfer of the heated by the combustion compressor cooling air K in the turbine. Since this mechanical container is located near the flame, it is exposed to temperatures that exceed even the melting temperature of superalloys. Therefore, in order to prevent the combustion chambers from melting, they are often equipped with complex double-walled cooling systems and cooling fins between the walls.
  • a combustion chamber for a diffusion flame, which has a double wall, is, for example, in WO 99/17057 A1 described.
  • the first object is achieved by a gas turbine combustor according to claim 1, the second object by a gas turbine according to claim 12.
  • the appended claims contain advantageous embodiments of the invention.
  • a gas turbine combustor according to the invention has a substantially rotationally symmetrical cross-section and has at least one axial section which has an inner wall with an outer side and an outer wall with an inner side facing the inner wall and spaced from the inner side.
  • a gap forming at least one cooling fluid channel.
  • the outer wall to the space leading inlet openings for a cooling fluid, preferably compressor cooling air K on.
  • a porous, metallic structure is provided at least partially in the intermediate space.
  • the porous metallic structure may vary in its radial height, e.g. form a continuous layer between inner and outer wall, i. fill the height completely or fill the height only partially.
  • the use of a metallic porous structure in the space between outer wall and inner wall improves the heat dissipation in two ways. First, by increasing the heat transfer, since no developed thermal boundary layer can build.
  • the thermal boundary layer is the area of a fluid which is influenced by a heat flow from or into a wall when the wall has a different temperature than the fluid.
  • another fluid can also generate a thermal boundary layer.
  • the thermal boundary layer is bounded on the one hand by the wall, on the other hand by an imaginary surface, at which the temperature no longer changes in the direction of the interior of the fluid.
  • the thickness of the thermal boundary layer increases in the direction of the flow, since, depending on the wall temperature, heat is supplied or withdrawn from the fluid. If the fluid flows in a pipe or a channel, the thermal boundary layers can grow together from both sides after a certain distance in the middle. From there, the area-related heat transfer performance decreases, since the temperature difference between wall and core flow also decreases.
  • the heat transfer performance So can not be increased arbitrarily by extending the flow path.
  • the porous structure also allows for more efficient cooling than film cooling or impingement cooling alone because the porous structure is highly surface enlarging. It allows a kind of transpiration cooling.
  • the metallic structure comprises an Invar alloy.
  • This alloy has abnormally small or sometimes negative coefficients of thermal expansion in certain temperature ranges.
  • deformation can be e.g. the buckling combustion chamber walls are counteracted.
  • the height of the inner wall can be arbitrary, that is also such that it adjoins the outer wall on the combustion chamber side.
  • the inner wall is preferably provided with a protective layer (TBC) on the combustion chamber side. This causes a further hot gas protection.
  • TBC protective layer
  • the inner wall has a downstream end at which the gap between the outer side of the inner wall and the inner side of the outer wall is open towards the interior of the combustion chamber.
  • the cooling air / cooling fluid used can be blown out (film cooling) after passing through the porous structure in the combustion chamber and participate in the combustion.
  • the outer wall and the inner wall are integrally formed in the axial direction of the gas turbine combustor.
  • the production can be significantly simplified.
  • a ceramic thermal barrier coating or another TBC on the hot gas side ie the combustion chamber inside it is possible to dispense with the blowing out of the cooling air / cooling fluid as far as possible or even completely.
  • the outer wall has in the axial direction of the gas turbine combustor stages and a number of axially disposed behind inner walls, the inner walls are annular and the diameters of axially successively arranged annular inner walls increase and wherein adjacent annular inner walls partially are pushed into each other.
  • each of the outer walls of the inner walls which are partially pushed into one another, may be fastened to a fastening section of the outer wall in their section surrounding the inner of the inner walls pushed one inside the other.
  • the inlet openings of the outer wall then adjoin these fastening sections.
  • Each intermediate space formed between an inner wall and the outer wall can then be supplied with cooling fluid individually.
  • each of the axially arranged in succession inner walls have a downstream end on which the existing between the outside of the respective inner wall and the inside of the outer wall gap to the combustion chamber interior is open. In this way, a further inner wall arranged in the axial direction behind an inner wall is further cooled by means of film cooling by the cooling fluid entering the combustion chamber, which flows along the inside of the following inner wall.
  • inlet openings are provided. Through these inlet openings, the cooling fluid / cooling air can flow into the gap and there through the porous structure, and thus efficiently cool the combustion chamber inner wall.
  • the cooling fluid / cooling air is discharged axially along the combustion chamber walls.
  • a specification of the flow direction that is, in the direction of the turbine or compressor is possible, or else a discharge of the cooling fluid in both directions.
  • the inlet openings are provided over the entire axial length of the outside. Air or cooling fluid is passed through these inlet openings in the space with the porous structure and flows through the porous structure.
  • the inlet openings can be adjusted to the places be that require particularly good or high cooling on the combustion chamber wall.
  • the inlet openings are preferably installed locally, so that the compressor cooling air K passing through the inlet openings is flowed into the intermediate space in front of the porous, metallic structure as seen in the flow direction.
  • a gas turbine according to the invention is equipped with at least one combustion chamber according to the invention.
  • a plurality of combustion chambers according to the invention for example six, eight or twelve combustion chambers, may be arranged around the rotor.
  • the advantages described with reference to the gas turbine combustor according to the invention also result in the gas turbine according to the invention. Reference is therefore made to the described advantages of the gas turbine combustor according to the invention.
  • the outer wall has passages.
  • the porous structure can also be used to damp the combustion chamber vibrations, e.g. by means of mounted in the wall cooling holes that can serve as a sound absorber.
  • FIG. 1 shows a gas turbine 1 in a longitudinal section.
  • This comprises a compressor section 3, a combustion chamber section 5 and a turbine section 7.
  • a shaft extends through all sections of the gas turbine 1.
  • the shaft 9 is provided with rings of compressor blades 11 and in the turbine section 7 with rings of turbine blades 13.
  • Wreaths of compressor vanes 15 are located in the compressor section 3 between the rotor blade rings and rings of turbine vanes 17 in the turbine section 7.
  • the vanes extend from the housing 19 of the gas turbine installation 1 essentially in the radial direction to the shaft.
  • FIG. 2 shows a combustion chamber 25 according to the invention a gas turbine 1 in a schematic sectional view.
  • the combustion chamber 25 includes a burner end 31 to which at least one burner 27 is disposed and through which both the fuel and compressor air are introduced into the combustion chamber.
  • the combustion chamber 25 comprises a turbine-side outlet end 33, through which the hot combustion exhaust gases exit the combustion chamber 25 in the direction of the turbine section 7.
  • the flame present in the combustion chamber 25 during operation of the gas turbine 1 results in very high temperatures in a section 35 of the combustion chamber which necessitate cooling of the combustion chamber wall, in particular when the flame is a diffusion flame.
  • the combustion chamber wall at least in this section 35, has a double-walled structure with an outer wall 37 and one or more inner walls 39A, 39B, 39C. Between the inner walls 39A, 39B, 39C and the outer wall 37 there are intermediate spaces 41A, 41B, 41C which form cooling fluid passages for a cooling fluid, in the present embodiment compressor air.
  • the inner walls 39 can ( Fig. 2 ) each have a mounting portion 45, in which they are attached to a mounting portion 46 of the outer wall 37.
  • the inner walls 39 have slightly different radii, wherein the radii in the flow direction 47 of the combustion gases increase.
  • the fastening portions 45 remote from the ends 40 of the inner walls 39 are inserted into a part in the downstream adjacent inner wall 39.
  • a distance between the outside of the inner inner wall (eg 39A) and the inner side of the outer Innwand (eg 39B) or the outer wall 37 remains such that on the outflow side an annular opening 42 open towards the combustion chamber interior is formed.
  • the intermediate spaces 41 are now at least partially filled with a porous, metallic structure 30 according to the invention. This can also vary in the radial height.
  • the outer wall 37 has, in the vicinity of the fixing portions 46 to which the inner walls 39 are fixed with their fixing portions 45, through holes 49 serving as inlet openings for compressor air into the spaces 41.
  • the compressor air thus flows axially through the porous structure 30 to cool it as well as the inner wall 39 and outer wall 37. Finally, the compressor air flows through the annular opening 42 into the combustion chamber interior. Due to the significantly increased surface to be cooled, a more efficient cooling now takes place.
  • porous structure 30 in the space between outer 37 and inner wall 39 improves the heat dissipation in two ways. On the one hand by increasing the heat transfer surfaces, as well as by increasing the heat transfer, since no developed boundary layer can build.
  • the porous structure 30 can also serve to dampen combustion oscillations. For example, the texture, e.g. large or small pores on different areas of the walls to be matched to the required damping.
  • Figure 3 shows in detail such a porous structure 30. It can be dispensed with in this further embodiment, the mounting portion 45,46.
  • the passage holes 49 for the cooling air supply are arranged here at individual positions. However, the individual positions can be arranged over the entire length of the outer wall 37.
  • Figure 4 shows a combustion chamber comprising only a single segment (designed as a one-piece cylinder). Due to this improved convective cooling can be dispensed with blowing the cooling air into the combustion chamber. It is thus an open or closed cooling possible. The porosity significantly increases the coolable surface. Blowing the compressor cooling air into the combustion chamber is therefore no longer necessary. This facilitates the manufacture of the combustion chamber relative to the combustion chamber in the prior art. For example, if passages (not shown) are mounted in the inner wall 39, e.g. Holes so here, too, the porous structure 30 can be used for targeted damping of combustion chamber vibrations.
  • the flow direction of the cooling air can be as in Fig. 3 be predetermined (here in the direction of the turbine or possibly in the direction of the compressor, with appropriate adjustment of the annular opening 42). But it is also conceivable axial discharge of the cooling air in both directions.
  • the invention presented here can be applied to any burner which has a combustion chamber, in particular also burners with a simple, single-walled cylindrical and non-cylindrical outer wall.
  • porous metallic structures 30 can be mounted on the outside of the burner to also improve heat dissipation.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP09162297A 2009-06-09 2009-06-09 Chambre de combustion de turbine à gaz et turbine à gaz Withdrawn EP2261565A1 (fr)

Priority Applications (1)

Application Number Priority Date Filing Date Title
EP09162297A EP2261565A1 (fr) 2009-06-09 2009-06-09 Chambre de combustion de turbine à gaz et turbine à gaz

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP09162297A EP2261565A1 (fr) 2009-06-09 2009-06-09 Chambre de combustion de turbine à gaz et turbine à gaz

Publications (1)

Publication Number Publication Date
EP2261565A1 true EP2261565A1 (fr) 2010-12-15

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EP09162297A Withdrawn EP2261565A1 (fr) 2009-06-09 2009-06-09 Chambre de combustion de turbine à gaz et turbine à gaz

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EP (1) EP2261565A1 (fr)

Citations (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE1476727A1 (de) * 1964-10-20 1969-09-25 Bristol Siddeley Engines Ltd Kanalanordnung zur Kuehlung von Wandungen,die ein heisses stroemendes Medium von einem kuehleren stroemenden Medium hoeheren Druckes trennen
GB2034874A (en) * 1978-11-03 1980-06-11 Gen Electric Gas turbine engine combustor
US4262487A (en) * 1979-02-01 1981-04-21 Westinghouse Electric Corp. Double wall combustion chamber for a combustion turbine
US4838031A (en) * 1987-08-06 1989-06-13 Avco Corporation Internally cooled combustion chamber liner
WO1997014875A1 (fr) 1995-10-17 1997-04-24 Westinghouse Electric Corporation Dispositif de combustion refroidi regeneratuer pour turbine a gaz
WO1999017057A1 (fr) 1997-09-30 1999-04-08 Siemens Westinghouse Power Corporation CHAMBRE DE COMBUSTION A TRES FAIBLE EMISSION DE NO¿x?
EP1098141A1 (fr) * 1999-11-06 2001-05-09 Rolls-Royce Plc Eléments de paroi pour turbomachine
US6495207B1 (en) * 2001-12-21 2002-12-17 Pratt & Whitney Canada Corp. Method of manufacturing a composite wall
EP1512911A1 (fr) * 2003-09-04 2005-03-09 Rolls-Royce Deutschland Ltd & Co KG Arrangement pour refroidir des éléments soumis à de fortes contraintes thermiques
EP1712739A1 (fr) * 2005-04-12 2006-10-18 Siemens Aktiengesellschaft Elément avec trou de refroidissement par film

Patent Citations (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
DE1476727A1 (de) * 1964-10-20 1969-09-25 Bristol Siddeley Engines Ltd Kanalanordnung zur Kuehlung von Wandungen,die ein heisses stroemendes Medium von einem kuehleren stroemenden Medium hoeheren Druckes trennen
GB2034874A (en) * 1978-11-03 1980-06-11 Gen Electric Gas turbine engine combustor
US4262487A (en) * 1979-02-01 1981-04-21 Westinghouse Electric Corp. Double wall combustion chamber for a combustion turbine
US4838031A (en) * 1987-08-06 1989-06-13 Avco Corporation Internally cooled combustion chamber liner
WO1997014875A1 (fr) 1995-10-17 1997-04-24 Westinghouse Electric Corporation Dispositif de combustion refroidi regeneratuer pour turbine a gaz
WO1999017057A1 (fr) 1997-09-30 1999-04-08 Siemens Westinghouse Power Corporation CHAMBRE DE COMBUSTION A TRES FAIBLE EMISSION DE NO¿x?
EP1098141A1 (fr) * 1999-11-06 2001-05-09 Rolls-Royce Plc Eléments de paroi pour turbomachine
US6495207B1 (en) * 2001-12-21 2002-12-17 Pratt & Whitney Canada Corp. Method of manufacturing a composite wall
EP1512911A1 (fr) * 2003-09-04 2005-03-09 Rolls-Royce Deutschland Ltd & Co KG Arrangement pour refroidir des éléments soumis à de fortes contraintes thermiques
EP1712739A1 (fr) * 2005-04-12 2006-10-18 Siemens Aktiengesellschaft Elément avec trou de refroidissement par film

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