EP3726008B1 - Conduit de transition pour un ensemble de turbine à gaz et ensemble de turbine à gaz comportant ce conduit de transition - Google Patents

Conduit de transition pour un ensemble de turbine à gaz et ensemble de turbine à gaz comportant ce conduit de transition Download PDF

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Publication number
EP3726008B1
EP3726008B1 EP19170045.9A EP19170045A EP3726008B1 EP 3726008 B1 EP3726008 B1 EP 3726008B1 EP 19170045 A EP19170045 A EP 19170045A EP 3726008 B1 EP3726008 B1 EP 3726008B1
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EP
European Patent Office
Prior art keywords
transition duct
combustor
groove
gas turbine
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.)
Active
Application number
EP19170045.9A
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German (de)
English (en)
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EP3726008A1 (fr
Inventor
Michael Thomas Maurer
Hans-Christian Mathews
Patrik Meng
Klaus Michael DUESING
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Ansaldo Energia Switzerland AG
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Ansaldo Energia Switzerland AG
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Publication date
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Priority to EP19170045.9A priority Critical patent/EP3726008B1/fr
Priority to CN202010304341.3A priority patent/CN111829013B/zh
Publication of EP3726008A1 publication Critical patent/EP3726008A1/fr
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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
    • 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
    • F23R3/60—Support structures; Attaching or mounting means
    • 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
    • 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
    • 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
    • F23R3/46—Combustion chambers comprising an annular arrangement of several essentially tubular flame tubes within a common annular casing or within individual casings
    • 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
    • F05D2230/00—Manufacture
    • F05D2230/30—Manufacture with deposition of material
    • F05D2230/31—Layer deposition
    • 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
    • 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
    • F05D2260/94—Functionality given by mechanical stress related aspects such as low cycle fatigue [LCF] of high cycle fatigue [HCF]
    • F05D2260/941—Functionality given by mechanical stress related aspects such as low cycle fatigue [LCF] of high cycle fatigue [HCF] particularly aimed at mechanical or thermal stress reduction
    • 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
    • F23R2900/00017—Assembling combustion chamber liners or subparts
    • 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
    • F23R2900/00018—Manufacturing combustion chamber liners or subparts
    • 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

Definitions

  • the present invention relates to the technical filed of the gas turbine assemblies for power plants.
  • the present invention refers to a particular component of a gas turbine assembly, i.e. a duct (called in this field as “transition” duct) configured for guiding the hot gas leaving a combustion chamber towards a turbine inlet.
  • a duct called in this field as “transition” duct
  • the present invention refer to the tubular wall structure of the above transition duct (called in this field as "liner”) .
  • each can-combustor comprises an upstream first combustor (called in this field as “premix” combustor) configured for receiving compressed air and mixing this air with fuel, a downstream second combustor (called in this field as “reheat” combustor) configured for receiving the hot gas leaving the first combustor and adding fuel into this hot gas for performing a self/spontaneous ignition and a transition duct (as foregoing mentioned) for guiding the hot gas leaving the reheat combustor to the turbine inlet.
  • premix premix
  • reheat downstream second combustor
  • the liner of the present invention may be defined as a "sequential" liner.
  • the present invention is not limited to the above filed of combustors having a reheat configuration. Indeed, the present invention may also be applied without any modification to a non-reheat combustor.
  • a gas turbine assembly for power plants comprises a rotor having an axis (i.e. the gas turbine axis), a compressor, a combustor unit and at least a turbine.
  • the compressor is configured for compressing air supplied at a compressor inlet.
  • the compressed air leaving the compressor flows into a plenum and from there into the combustor unit.
  • the combustor unit comprises a plurality of burners configured for injecting fuel in the compressed air flow.
  • the mixture of fuel and compressed air flows into a combustion chamber where this mixture is combusted.
  • the resulting hot gas leaves the combustion chamber and is expanded in the turbine performing work on the rotor.
  • the turbine comprises a plurality of stages, or rows, of rotor blades that are interposed by a plurality of stages, or rows, of stator vanes.
  • the rotor blades are connected to the rotor whereas the stator vanes are connected to a vane carrier that is a concentric casing surrounding the turbine unit.
  • a high turbine inlet temperature is required.
  • this high temperature involves an undesired high NOx emission level.
  • the so called “sequential" gas turbines is particularly suitable.
  • a sequential gas turbine comprises two combustors or combustion stages in series wherein each combustor is provided with a plurality of burners and with at least a relative combustion chamber. Following the main gas flow direction, usually the upstream or first combustor comprises a plurality of so-called "premix" burners.
  • each burner of the first combustor is configured not only for injecting the fuel directly in the compressed air (for instance with a so called diffusion flame) but also for mixing (with a swirl) the compressed air and the fuel before injecting the mixture into the combustion chamber.
  • the downstream or second combustor is called “reheat” or “sequential” combustor and it is fed by the hot gas leaving the first combustor.
  • the reheat combustor is provided with a plurality of reheat burners configured for injecting fuel in the hot gas coming from the first combustor. Due to the high gas temperature, the operating conditions downstream the reheat burners allow a self/spontaneous ignition of the fuel/air mixture.
  • each can-combustor comprises a premix (first stage) and a reheat (second stage) combustor arranged directly one downstream the other inside the common can shaped casing ending with a tubular element (called in this field as "transition" duct) configured for guiding the hot gas leaving the reheat (second stage) combustor toward the turbine inlet.
  • the transition duct requires cooling to avoid damages caused by overheating and in order to increase lifetime.
  • a fraction of the total airflow is usually taken from the compressor (i.e. before the combustor) and used as convective cooling air acting on the outer surface of the transition duct.
  • Today is present the need to improve the cooling of the transition duct in order to allow further raising the firing temperature inside the transition duct itself.
  • a common solution of this problem is to use a higher amount of cooling air.
  • a higher air consumption used for cooling purpose reduces the efficiency of the gas turbine.
  • a solution developed by the Applicant consists in providing the tubular wall of the transition duct (i.e.
  • the tubular wall of the transition duct comprises an inner wall portion (called in this field as “hot shell”), an outer wall portion (called in this field as “cold shell”) and a plurality of ribs as partitions of adjacent cooling channels and connecting the inner wall portion and the outer wall portion.
  • the ribs may be defined as radial ribs considering the cross section of the transition duct even if not the entire transition duct may disclose a circular cross section.
  • ribs act as a stiff structure, indeed for this reason it is possible to realize a thin hot shell.
  • these radial ribs involves a higher creep resistance with respect to a single layer design.
  • This sequential liner (having cooling channels extending through the tubular wall divided by radial ribs) works well as long as the thermal barrier coating provided on the hot shell is not lost during the operation of the gas turbine. In case of loss of this thermal barrier coating (or in case of it is required a very high temperature difference between the hot and could shell), due to the presence of these rigid connections (i.e. the radial ribs) a static stress level increases in the wall structure.
  • EP2206886 and US2006/0185345 disclose transition ducts for gas turbines.
  • EP2206886 discloses a transition duct according to the preamble of claim 1.
  • a transition duct as set out in claim 1 and a gas turbine for power plant as set out in claim 10 are provided.
  • a primary object of the present invention is to provide a transition duct for a gas turbine configured for guiding the hot gas from the combustor to the turbine wherein this transition duct is also suitable for overcoming the prior art limits foregoing mentioned.
  • a transition duct is provided wherein this transition duct comprises a tubular wall provided with a plurality of cooling channels and wherein this transition duct allows to reach an higher thermal gradient between the hot shell and the cold shell forming the wall.
  • this transition duct comprises:
  • the connecting structure is a flexible connecting structure configured for allowing a circumferential movement of the outer surface (or outer wall portion) whit respect to the inner surface (or inner wall portion).
  • the circumferential direction is defined with respect to the longitudinal axis of the duct (main hot gas flow direction) that usually comprises at least a portion having a circular cross section. Due to the fact that the outer surface can grow along the circumferential direction (instead of the rigid configuration of the prior art practice), it is possible to increase the thermal mismatch between the outer surface and the inner surface without the risk of damaging the wall structure.
  • the duct is provided with at least a groove extending from the cold shell toward the hot shell and/or with an elongated portion of the cold shell directed inside a channel toward the hot shell or directed outside the channel.
  • the elongated portion may be preferably V-shaped, U-shaped or half-moon shaped or any similar shape allowing a circumferential movement of the cold shell.
  • the groove and/or the elongated portion may also be provided with a passing hole for allowing the cooling air to enter a channel.
  • the transition duct of the present invention is a single piece made of an additive manufacturing method, for instance a selective laser melting process.
  • the connecting structure may also be configured for controlling another movement of the wall structure.
  • the connecting structure may also be configured for avoiding any radial movements of the inner surface or of the wall structure in general. Therefore, according to this embodiment the radial movement of the structure is prohibited/inhibited also in case of creep.
  • the radial direction is defined with respect to the longitudinal axis of the duct (or the main hot gas flow direction) that usually comprises at least a portion having a circular cross section.
  • this additional feature of the connecting structure has been above defined in a functional manner because the skill person may easily provide a plurality of different embodiments allowing this added result. In any case, in the description of the drawings some different embodiments provided with this feature will be described.
  • This feature is implemented according to the present invention by providing the duct with a circumferential indentation housed in a circumferential seat or extending as a lip above a portion of the could shell.
  • the new connecting structure of the present invention does not affect the shape, geometry or configuration of inner surface of the duct. Therefore, the implementation of the present invention does not interfere with the thermal barrier coating applied on the inner surface.
  • the present invention also refers to a gas turbine for power plant wherein this gas turbine has an axis and comprises:
  • the combustor sector of the present invention comprises a plurality can combustors, wherein each can combustor may comprise a single combustion stage or in series a first combustor a second combustor.
  • Each can combustor comprise also a transition duct that is realized according to the enclosed claims for guiding the hot gas from the combustor to the turbine.
  • figure 1 is a side elevation view, cut along an axial, longitudinal plane, of a gas turbine assembly that can be provided with the transition duct according to the present invention.
  • FIG 1 discloses a simplified view of a gas turbine assembly, designated as whole with reference 1.
  • the gas turbine assembly 1 comprises a compressor 2, a combustor assembly 3 and a turbine 4.
  • the compressor 2 and the turbine 5 extend along a main axis A.
  • the combustor assembly 3 disclosed in the example of figure 1 is a can combustor 5 that may be a sequential combustor or a single-stage combustor. Therefore, in this embodiment, the combustor assembly 3 comprises a plurality of sequential can combustors 5 circumferentially arranged about the main axis A.
  • the compressor 2 of the gas turbine engine 1 provides a compressed airflow, which is added with fuel and burned in the can combustors 5. Part of the airflow delivered by the compressor 2 is also supplied to the combustor assembly 3 and to the turbine section 4 for the purpose of cooling.
  • figure 2 discloses a can combustor 5 of the gas turbine assembly 1 of figure 1 .
  • the can combustors 5 disclosed in figure 2 comprises a first-stage combustor 6 and a second-stage combustor 7 and a transition duct 8, sequentially arranged and defining a hot gas path.
  • the first-stage combustor 6 comprises a first-stage burner unit 9 and a first-stage combustion chamber 10.
  • the second-stage combustor 7 is arranged downstream of the first-stage combustor 6 and comprises a second-stage burner unit 17 and a second-stage combustion chamber 18.
  • the second-stage combustor 7 is furthermore coupled to the turbine 4, here not shown, through the transition duct 8.
  • the second-stage combustion chamber 12 extends along an axial direction downstream of the first-stage combustor 6.
  • the second-stage combustion chamber 12 comprises an outer liner 13 and an inner liner 14 wherein the outer liner 13 surrounds the inner liner 14 at a distance therefrom, so that a convective cooling channel 15 is defined between the outer liner 13 and the inner liner 14.
  • the disclosed transition duct 8 is the component of the hot gas path suffering the most severe thermal stress. Therefore, for cooling purpose the transition duct 8 comprises a tubular wall structure 17 provided with a plurality of cooling channels 16 fed by cooling air.
  • the tubular wall structure 17 has an upstream end 18 having a circular cross section, a downstream end 19 having a substantially rectangular cross section, an inner surface 20 and an outer surface 21.
  • the upstream end 18 is joined to the second-stage combustion chamber 12, whereas the downstream end 19 faces the turbine 4.
  • the inner surface 20 delimits a hot gas flow volume through which hot gas flows to reach the turbine 4.
  • the inner surface 20 is therefore directly exposed to hot gas flowing through the hot gas path.
  • the inner surface may be coated at least in part by a thermal barrier coating.
  • the cooling channels 16 extend through the tubular wall structure 17 between the upstream end 18 and the downstream end 19 and are uniformly distributed in a circumferential direction of the tubular wall structure 17. At the upstream end 18, the cooling channels 16 are in fluid communication with the convective cooling channel 15 of the second-stage combustion chamber 12. In this embodiment, the cooling channels 16 extend in an axial longitudinal direction of the tubular wall structure 17.
  • FIG 5A is an enlarged view of the portion labelled with the reference V in in figure 4 .
  • This figure 5A discloses a wall 17 according to the prior art practice (rigid connection).
  • the cooling channels 16 separate different portions of the tubular wall structure 17.
  • an inner wall portion 22 or hot shell of the tubular wall structure 17 that is the portion of the wall between the inner surface 20 and the cooling channels 16.
  • an outer wall portion 23 or cold shell of the tubular wall structure 17 may be defined as the portion of the wall between the outer surface 21 and the cooling channels 16.
  • Adjacent cooling channels 16 are separated by diaphragms or partitions in form or radial ribs 24 extending between the inner wall portion 22 and to the outer wall portion 23 of the tubular wall structure 17.
  • the temperature of the hot shell increases by some 100K more with respect to the temperature of the cold shell (outer wall portion 23).
  • any further increase of heat load or the loss of the thermal barrier coating applied on the hot shell lead to a parallel increase on this temperature difference.
  • the hot shell wants to circumferentially expand more than the cold shell, the cold shell has to resists to an high static stress load acting along the circumferential direction.
  • the direction references A, R e C are disclosed wherein the direction A is the axial direction (main hot gas flow direction), the direction C is the circumferential direction and the direction R is the radial direction.
  • the diagram of figure 5B schematically discloses the static stress load acting on the wall structure.
  • the proposed solution is to provide the wall 17 with a flexibility feature configured for allowing the cold shell 23 to grow in the circumferential direction. In this way the stress level acting on the cold shell due to the thermal mismatch is reduced and compensate by this circumferential movement.
  • the following description of the figures 6-11 will refer to some particular embodiments of the invention, i.e. embodiments wherein the rigid prior art connection between the hot shell and the cold shell (radial ribs) has been modified to allow the cold shell to freely grow in the circumferential direction. In any case, in all disclosed embodiments the radial movement of the hot shell is prohibited/inhibited also in case of creep.
  • the transition duct wall structure 17 comprises at least a radial groove or slot 25 (axially extending at least in part along the duct) realized inside a rib 24 and extending from the could shell 23 towards the hot shell 22. Therefore, at this radial groove 25 the thickness of the transition duct wall structure 17 is reduce to the sole hot shell 22. In other words, at this radial groove or slot 25 along the circumferential direction the integrity of the transition duct wall structure 17 is performed only at a point the hot shell 22 acting as a hinge 28 for the structure.
  • the cold shell 23 discloses two separated facing edges allowing the cold shell to move along the circumferential direction C.
  • a middle portion of groove or slot 25 along the radial direction is provided at one side with a circumferential indentation 26 and at the opposite side with a circumferential seat 27 housing at least in part the circumferential indentation 26.
  • FIG. 7 is a schematic view of an alternative embodiment of a transition duct wall structure according to the present invention.
  • This embodiment differs from the previous for the position of the circumferential indentation 26 along the radial direction of the groove or slot 25.
  • the indentation 26 is extending outside the groove or slot 25, in particular outwardly the cold shell 23 radially covering (spaced) the groove or slot 25. Also in this case the indentation 26 allows to prevent any radial movement of the hot shell.
  • the transition duct wall structure 17 comprises at least a V shaped groove or lowered portion 29 of the could shell 23 at a channel 16 wherein this V shaped portion is directed towards the hot shell 22.
  • a bridge 30 may be provided for connecting the apex of the V shaped portion 29 to the hot shell 22.
  • the stiffness of the cold shell is reduced to allow the cold shell itself to move along the circumferential direction.
  • FIG. 9 is a schematic view of an alternative embodiment of a transition duct wall structure according to the present invention.
  • This embodiment differs from the previous for the fact that the V shaped portion 29' is not extending towards the hot shell 23 but outside the channel 16. Also in this case at the V shaped portion 29' the stiffness of the cold shell 23 is reduced to allow the cold shell itself to move along the circumferential direction.
  • FIG. 10 is a schematic view of an alternative embodiment of a transition duct wall structure according to the present invention.
  • This embodiment comprises all features of the embodiment of figure 6 and moreover it discloses at least a cooling hole 30 connecting the groove 25 to an adjacent channel 16.
  • the cooling hole 30 is realized at the end of the groove 25 opposite to the cold shell 23.
  • the arrow F in figure 10 represent the cooling flow passing inside the groove 25 and the hole 30 for entering a channel 16.
  • FIG. 11 is a schematic view of an alternative embodiment of a transition duct wall structure according to the present invention.
  • This embodiment is quite similar to the embodiment disclosed in figure 7 .
  • the embodiment of figure 11 comprises a cooling hole 30 for connecting the groove 25 to a channel 16 (as the previous example) and it is not provided with the indentation 26 covering the groove 25. Indeed, this indentation 26 may limit the air flow entering the groove 25.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Claims (10)

  1. Conduit de transition (8) pour un ensemble de turbine à gaz (1) et configuré pour guider les gaz chauds d'une chambre de combustion vers une turbine, le conduit de transition (8) comprenant :
    - une structure de paroi tubulaire (17) pourvue d'une surface intérieure (20) en contact avec un gaz chaud et d'une surface extérieure (21) en contact avec un air de refroidissement ;
    - une structure de connexion s'étendant entre la surface intérieure (20) et la surface extérieure (21) et configurée pour définir une pluralité de canaux de refroidissement (16) s'étendant à travers la structure de paroi tubulaire (17) ;
    la structure de connexion est une structure de connexion flexible configurée pour permettre un mouvement circonférentiel de la surface extérieure (20) par rapport à la surface intérieure (21) et comprenant une pluralité de nervures (24) s'étendant de la surface extérieure (21) à la surface intérieure (20) pour séparer des canaux adjacents (16) ;
    caractérisé en ce que
    le conduit de transition (8) comprend au moins une rainure (25) s'étendant à l'intérieur d'une nervure (24) depuis la surface extérieure (21) vers la surface intérieure (20).
  2. Conduit de transition (8) selon la revendication 1, dans lequel le conduit de transition (8) est une pièce unique par un procédé de fabrication additive.
  3. Conduit de transition (8) selon la revendication 1, dans lequel une partie médiane de la rainure (25) est pourvue d'un côté d'une indentation circonférentielle (26) et du côté opposé d'un logement circonférentiel (27) abritant au moins en partie l'indentation (26).
  4. Conduit de transition (8) selon la revendication 1, dans lequel la surface extérieure (21) est munie d'une partie de lèvre recouvrant vers l'extérieur la rainure (25).
  5. Conduit de transition (8) selon l'une quelconque des revendications précédentes, dans lequel la surface extérieure (21) comprend une partie allongée (29') s'étendant à l'intérieur d'un canal (16) vers la surface intérieure (20).
  6. Conduit de transition (8) selon la revendication 5, dans lequel un pont (30) reliant la partie allongée (29') et la surface intérieure (20) est prévu.
  7. . Conduit de transition (8) selon la revendication 5, dans lequel la surface extérieure (21) comprend une partie allongée (29') s'étendant à l'extérieur d'un canal (16).
  8. Conduit de transition (8) selon l'une quelconque des revendications précédentes, dans lequel la rainure (25) ou la partie allongée comprend au moins un trou (30) reliant la rainure (25) ou la partie allongée à un canal (16).
  9. Conduit de transition (8) selon la revendication 8, dans lequel le trou (30) est prévu à l'extrémité de la rainure (25) opposée à la surface extérieure (21).
  10. Turbine à gaz pour centrale électrique ; la turbine à gaz (1) présentant un axe (A) et comprenant, en suivant le sens d'écoulement de gaz :
    - un compresseur (2) pour comprimer l'air,
    - un secteur de chambre de combustion (3) pour mélanger et brûler l'air comprimé avec au moins un combustible
    - une turbine (4) pour détendre le flux de gaz chauds brûlés quittant les chambres de combustion (3) ;
    dans lequel la chambre de combustion (4) comprend une pluralité de chambres de combustion (5), chaque chambre de combustion logeant une seule chambre de combustion ou en série une première chambre de combustion (6), une deuxième chambre de combustion (7) et un conduit de transition (8), le conduit de transition (8) étant réalisé selon l'une quelconque des revendications précédentes.
EP19170045.9A 2019-04-18 2019-04-18 Conduit de transition pour un ensemble de turbine à gaz et ensemble de turbine à gaz comportant ce conduit de transition Active EP3726008B1 (fr)

Priority Applications (2)

Application Number Priority Date Filing Date Title
EP19170045.9A EP3726008B1 (fr) 2019-04-18 2019-04-18 Conduit de transition pour un ensemble de turbine à gaz et ensemble de turbine à gaz comportant ce conduit de transition
CN202010304341.3A CN111829013B (zh) 2019-04-18 2020-04-17 用于燃气涡轮组件的过渡管及包括该过渡管的燃气涡轮组件

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Application Number Priority Date Filing Date Title
EP19170045.9A EP3726008B1 (fr) 2019-04-18 2019-04-18 Conduit de transition pour un ensemble de turbine à gaz et ensemble de turbine à gaz comportant ce conduit de transition

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EP3726008A1 EP3726008A1 (fr) 2020-10-21
EP3726008B1 true EP3726008B1 (fr) 2022-05-18

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CN111829013B (zh) 2023-09-26
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