EP3078914A1 - Chambre de combustion annulaire pour un moteur à turbine à gaz - Google Patents

Chambre de combustion annulaire pour un moteur à turbine à gaz Download PDF

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Publication number
EP3078914A1
EP3078914A1 EP15162999.5A EP15162999A EP3078914A1 EP 3078914 A1 EP3078914 A1 EP 3078914A1 EP 15162999 A EP15162999 A EP 15162999A EP 3078914 A1 EP3078914 A1 EP 3078914A1
Authority
EP
European Patent Office
Prior art keywords
annular
gas turbine
combustor
connection element
turbine engine
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Withdrawn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
EP15162999.5A
Other languages
German (de)
English (en)
Inventor
Patrik Rasmusson
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 EP15162999.5A priority Critical patent/EP3078914A1/fr
Priority to PCT/EP2016/056793 priority patent/WO2016162239A1/fr
Priority to EP16713399.0A priority patent/EP3256782A1/fr
Priority to US15/563,687 priority patent/US20180073738A1/en
Publication of EP3078914A1 publication Critical patent/EP3078914A1/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Withdrawn legal-status Critical Current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
    • F23R3/005—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
    • F23R3/60—Support structures; Attaching or mounting means
    • 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
    • F23R2900/00012—Details of sealing devices

Definitions

  • the present invention relates to an annular combustor for a gas turbine engine and to a gas turbine engine including such a combustor.
  • Annular combustors are well known in the field of gas turbine engines.
  • An annular combustor is normally included in a gas turbine engine 1 comprising, arranged in flow series: a compressor section, a burner, the annular combustor and a gas turbine section.
  • air is compressed by the compressor section and delivered to the combustion section, including the burner and the annular combustor.
  • the compressed air exiting from the compressor enters the burner, where is mixed with a gaseous or liquid fuel.
  • the air/fuel mixture is then burned and the combustion gas from the combustion is channelled through the combustor to the gas turbine section, for transforming the energy from the operative gas into working power.
  • the annular combustor for the above described application normally extends axially between a first axial burner end, close to the burner and a second axial outlet end, adjacent to the inlet of the gas turbine section. Further the annular combustor radially extends between an annular inner liner and an annular outer liner. At first axial burner end, the annular combustor comprises an annular backwall for connecting the annular inner liner and the annular outer liner. The backwall comprises at least a hole for coupling at least a respective burner to the annular combustor.
  • the annular combustor described above is normally manufactured in one piece including at least the annular inner liner, the annular outer liner and the backwall. Such unitary piece is then attached to the gas turbine engine by means of one or more connections provided between the combustor and a casing of the gas turbine engine. Such a connection is typically remote from the interface between the outlet of the combustor and the inlet of the gas turbine, therefore a gap is necessary between the combustor and the gas turbine for allowing thermal expansions.
  • annular combustor for a gas turbine is provided in accordance with the independent claim.
  • the dependent claims describe advantageous developments and modifications of the invention.
  • annular combustor for a gas turbine engine, the annular combustor axially extending between a first axial burner end and a second axial outlet end, the annular combustor radially extending between an annular inner liner and an annular outer liner, the annular combustor being at least an assembly of:
  • the segmentation of the annular combustor allows reducing at the minimum or eliminating the gap between the outlet of the combustor and the inlet of the gas turbine. As a result, hot gas ingestion is eliminated or effectively reduced, thus producing a much more robust design, which would be also easier and therefore cheaper to repair.
  • the first or the second part further includes a backwall for connecting the annular inner liner and the annular outer liner at the first axial burner end, the backwall comprising at least an hole for coupling a burner to the annular combustor.
  • the annular combustor is an assembly of the first part, the second part, and at least a third part including:
  • the annular combustor of the present invention is made part of two parts, one including the backwall, or is made of three parts, respectively including inner liner, outer liner and backwall.
  • these variants give the possibility to adapt the combustor design of the present invention to the design of different gas turbine engines, for example gas turbine engines having different overall dimensions.
  • a sealing is provided between the backwall and at least one of the inner liner and the annular outer liner.
  • the sealing may comprise at least a finger seal.
  • the sealing avoids leakages through the contacts between first, second and third part of the annular combustor at the backwall, i.e. where the first, second and third part contact each other.
  • the annular inner liner and/or the annular outer liner comprise a plurality of effusion holes for letting compressed air to enter the combustor through the annular inner liner and /or the outer liner, in order to cool the annular inner liner and /or the outer liner, respectively.
  • the annular inner liner and/or the annular outer liner comprises at least a cooling passage inside the liner.
  • the cooling passage may be provided between two panels of the annular inner liner and/or of the annular outer liner, bonded together.
  • the effusion holes or the cooling passages provide the necessary cooling to the walls of the inner and outer liners.
  • a gas turbine engine comprises a burner, a gas turbine and an annular combustor as above described, between the burner and the gas turbine.
  • a gas turbine comprises an inlet section and at least one connection element adjacent to the inlet section for coupling with an annular combustor as above described.
  • Figure 1 shows an example of a gas turbine engine 1 in a partial schematic sectional view.
  • the gas turbine engine 1 (not shown as a whole) comprises, in flow series, a compressor section 4 (not shown as a whole), a plurality of burners 2 (only one burner 2 shown in each of the section figures 1 to 4 ) an annular combustor 10 and a gas turbine 3, which are generally arranged in flow series within a casing 5.
  • the gas turbine engine 1 is generally arranged about a rotational axis X, which is the rotational axis for rotating components, in particular the compressor section 4 and the gas turbine 3.
  • the rotational axis X is also coincident with the axis of symmetry of the annular combustor 10, when the annular combustor 10 is assembled to the gas turbine engine 1.
  • air is compressed by the compressor section 4 and delivered to the combustion section, including the burner 2 and the annular combustor 10.
  • the compressed air exiting from the compressor 4 and flowing towards the combustion section is schematically represented in the attached figures by arrows A.
  • the compressed air enters the burner 2 where is mixed with a gaseous or liquid fuel.
  • the air/fuel mixture is then burned and the combustion gas from the combustion is channelled through the combustion chamber 10 to the gas turbine section 1, for transforming the energy from the operative gas into working power.
  • the combustion gas flows along the combustion chamber 10 along a main direction oriented from the burner 2 to the gas turbine 3.
  • the combustion section 4 and the burner 2 are not a specific object of the present invention and, therefore, in the following, they will not be described in further detail.
  • the annular combustor 10 extends axially between a first axial burner end 11 and a second axial outlet end 12 and radially between an annular inner liner 15 and an annular outer liner 16.
  • the annular combustor 10 further includes:
  • connection element 31 and of the second connection element 32 is adjacent to the outlet end 12 of the combustor 10 and is connectable to a respective gas turbine connection element 17, 18 on the gas turbine 3, respectively radially inner and outer.
  • the connection elements 17, 18 are adjacent to the inlet section 13 of the gas turbine 3.
  • the annular combustor 10 further comprises a third part 23, distinct from first part 21 and from the second part 22.
  • the third part 23 includes the backwall 25 and a third connection element 33 for connecting the third part 22 to the casing 5 of the gas turbine engine 1.
  • annular combustor 10 is an assembly of:
  • first, second and third part 21, 22, 23 and the gas turbine 3 and the casing 5, respectively is circumferentially distributed about the axis X. According to the different embodiments of the present invention, this may be obtained with circumferential elongated connection elements 31, 32, 33, 17, 18, 35.
  • each of the first part 21, second part 22 and third part 23 comprises a respective plurality of connection elements 31, 32, 33 to be coupled to a respective plurality of connection elements 17, 18 on the gas turbine 3 and of connection elements 35 on the casing 5.
  • the third part 23 is not present and the backwall 25 is comprised in the first part 21 and in the second part 22, respectively.
  • the annular combustor 10 of figure 2 is therefore an assembly of:
  • first and the second part 21, 22 and the casing and 5 the gas turbine 3, respectively is circumferentially distributed about the axis X. According to the different embodiments of the present invention, this may be obtained with circumferential elongated connection elements 31, 32, 18, 35.
  • each of the first part 21 and second part 22 comprises a respective plurality of connection elements 31, 32 to be coupled to a respective plurality of casing connection elements 35 and of outer connection elements 18 on the gas turbine 3.
  • the annular combustor 10 of figure 3 is instead an assembly of:
  • first and the second part 21, 22 and the gas turbine 3 and the casing 5, respectively is circumferentially distributed about the axis X. According to the different embodiments of the present invention, this may be obtained with circumferential elongated connection elements 31, 32, 17, 35.
  • each of the first part 21 and second part 22 comprises a respective plurality of connection elements 31, 32 to be coupled to a respective plurality of inner connection elements 17 on the gas turbine 3 and of casing connection elements 35.
  • the couplings between the first, second and third part 21, 22, 23 and the gas turbine 3 and/or the casing 5 are detachable and may performed by means of screw (or other threaded connection) and/or bolts. This allows connecting and disconnecting each part 21, 22, 23 of the annular combustor 10 independently from the others.
  • the mounting of the first, second and third part 21, 22, 23 in the gas turbine engine is made in such a way that the outlet end 12 of the annular combustor 10 is mounted adjacent to the inlet section 13 of the gas turbine 3. This avoids or limits hot gas leakages when hot gases from the annular combustor 10 enter the gas turbine 3.
  • the first, second and third part 21, 22, 23 contact each other along the edges of the backwall 25.
  • a sealing 40 is provided between the backwall 25 and at least one of the inner liner 15 and the annular outer liner 16.
  • the sealing 40 comprises an inner finger seal 41 between the backwall 25 and the inner liner 15 and/or an outer finger seal 42 between the backwall 25 and the outer liner 16.
  • both the inner and outer finger seals 41, 42 are present.
  • other sealing devices may be used between the parts 21, 22, 23 of the annular combustor 10, in order to avoid hot gas leakages between the backwall 25 and the inner liner 15 and/or between the backwall 25 and the outer liner 16.
  • the annular inner liner 15 and the annular outer liner 16 comprise a plurality of effusion holes 50 for letting compressed air (represented by arrows A) to enter the combustor 1) through the walls of the annular inner liner 15 and the outer liner 16, in order to cool the annular liners 15 and 16, respectively.
  • the embodiment of figure 4 is similar to the embodiment of figure 1 , i.e it comprises the first, second and third parts 21, 22, 23 of the annular combustor independently attached to the gas turbine 3 and to the casing 5.
  • the embodiment of figure 4 is different from the embodiment of figure 1 for the fact that the annular inner liner 15 and the annular outer liner 16 comprises one or more cooling passages 60 inside the respective liner 15, 16, for providing cooling by letting a flow of compressed air A enter the annular combustor 10 through the cooling passages 60.
  • Each cooling passage 60 is obtained inside the walls of the annular inner liner 15 and the annular outer liner 16 by means of at least two panels 61, 62, respectively internal (i.e.
  • Each cooling passage 60 comprises an inlet 63, through which compressed air A enters the passage 60, and an outlet 64, through which compressed air A exits the passage 60 to enter the annular combustor 10.

Landscapes

  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
EP15162999.5A 2015-04-09 2015-04-09 Chambre de combustion annulaire pour un moteur à turbine à gaz Withdrawn EP3078914A1 (fr)

Priority Applications (4)

Application Number Priority Date Filing Date Title
EP15162999.5A EP3078914A1 (fr) 2015-04-09 2015-04-09 Chambre de combustion annulaire pour un moteur à turbine à gaz
PCT/EP2016/056793 WO2016162239A1 (fr) 2015-04-09 2016-03-29 Chambre de combustion annulaire pour turbine à gaz
EP16713399.0A EP3256782A1 (fr) 2015-04-09 2016-03-29 Chambre de combustion annulaire pour turbine à gaz
US15/563,687 US20180073738A1 (en) 2015-04-09 2016-03-29 Annular combustor for a gas turbine engine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
EP15162999.5A EP3078914A1 (fr) 2015-04-09 2015-04-09 Chambre de combustion annulaire pour un moteur à turbine à gaz

Publications (1)

Publication Number Publication Date
EP3078914A1 true EP3078914A1 (fr) 2016-10-12

Family

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Family Applications (2)

Application Number Title Priority Date Filing Date
EP15162999.5A Withdrawn EP3078914A1 (fr) 2015-04-09 2015-04-09 Chambre de combustion annulaire pour un moteur à turbine à gaz
EP16713399.0A Withdrawn EP3256782A1 (fr) 2015-04-09 2016-03-29 Chambre de combustion annulaire pour turbine à gaz

Family Applications After (1)

Application Number Title Priority Date Filing Date
EP16713399.0A Withdrawn EP3256782A1 (fr) 2015-04-09 2016-03-29 Chambre de combustion annulaire pour turbine à gaz

Country Status (3)

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US (1) US20180073738A1 (fr)
EP (2) EP3078914A1 (fr)
WO (1) WO2016162239A1 (fr)

Cited By (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR3061761A1 (fr) * 2017-01-10 2018-07-13 Safran Aircraft Engines Chambre de combustion pour turbomachine

Families Citing this family (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
WO2015162795A1 (fr) * 2014-04-25 2015-10-29 三菱日立パワーシステムズ株式会社 Chambre de combustion de turbine à gaz et turbine à gaz équipée de ladite chambre de combustion
US10823418B2 (en) * 2017-03-02 2020-11-03 General Electric Company Gas turbine engine combustor comprising air inlet tubes arranged around the combustor
US11840032B2 (en) 2020-07-06 2023-12-12 Pratt & Whitney Canada Corp. Method of repairing a combustor liner of a gas turbine engine

Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1578474A (en) * 1976-06-21 1980-11-05 Gen Electric Combustor mounting arrangement

Family Cites Families (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5289677A (en) * 1992-12-16 1994-03-01 United Technologies Corporation Combined support and seal ring for a combustor

Patent Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB1578474A (en) * 1976-06-21 1980-11-05 Gen Electric Combustor mounting arrangement

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
FR3061761A1 (fr) * 2017-01-10 2018-07-13 Safran Aircraft Engines Chambre de combustion pour turbomachine
WO2018130765A1 (fr) * 2017-01-10 2018-07-19 Safran Aircraft Engines Chambre de combustion pour turbomachine
CN110168284A (zh) * 2017-01-10 2019-08-23 赛峰航空器发动机 涡轮发动机燃烧室
CN110168284B (zh) * 2017-01-10 2021-02-23 赛峰航空器发动机 涡轮发动机燃烧室
US11614234B2 (en) 2017-01-10 2023-03-28 Safran Aircraft Engines Turbine engine combustion chamber

Also Published As

Publication number Publication date
US20180073738A1 (en) 2018-03-15
WO2016162239A1 (fr) 2016-10-13
EP3256782A1 (fr) 2017-12-20

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