WO2009121820A1 - Combustor for a gas turbine - Google Patents

Combustor for a gas turbine Download PDF

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Publication number
WO2009121820A1
WO2009121820A1 PCT/EP2009/053701 EP2009053701W WO2009121820A1 WO 2009121820 A1 WO2009121820 A1 WO 2009121820A1 EP 2009053701 W EP2009053701 W EP 2009053701W WO 2009121820 A1 WO2009121820 A1 WO 2009121820A1
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WO
WIPO (PCT)
Prior art keywords
combustor
outer shell
burner
combustor liner
annular passage
Prior art date
Application number
PCT/EP2009/053701
Other languages
French (fr)
Inventor
Andreas Karlsson
Olle Lindman
Original Assignee
Siemens Aktiengesellschaft
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 Aktiengesellschaft filed Critical Siemens Aktiengesellschaft
Publication of WO2009121820A1 publication Critical patent/WO2009121820A1/en

Links

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
    • F23R3/54—Reverse-flow combustion chambers
    • 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
    • F23R3/50—Combustion chambers comprising an annular flame tube within an annular casing
    • 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 gas turbines, and in particular, to a serially cooled combustor for a gas turbine.
  • a gas turbine engine includes a compressor section, a combustor and a turbine section.
  • the compressor section delivers compressed air to the combustor which is used for combustion in a burner as well as for cooling the liner wall of the combustor, which must be maintained at an operating temperature meeting a durability requirement.
  • air from the compressor is first led through a cooling passage for cooling the liner wall of the combustor before flowing to the burner inlet.
  • This cooling passage is usually defined by the combustor liner and an outer shell, sometimes referred to as a cooling shield.
  • serial cooling leads to uneven inflow of air at the burner inlet causing combustor pulsations. This unevenness of air inflow to the burner is generally corrected by adjustment of the burner fuel profile to avoid such pulsations which results in unnecessarily high emissions.
  • the combustor liner is in direct contact with the hot combustion gases and is hence subjected to a much higher temperature than the outer shell.
  • This causes a thermal displacement between the hot combustor liner and the colder outer shell, as a result of which the connection of the hot combustor liner with the colder outer shell creates high thermal stresses in these components.
  • thermal displacements have been hitherto addressed by providing complicated moving parts for connecting the combustor liner and the outer shell.
  • the object of the present invention is to provide an improved combustor for a gas turbine.
  • the above object is achieved by a combustor for a gas turbine, comprising:
  • annular passage defined by a combustor liner and an outer shell, said annular passage forming a first portion of a flow-path for air, to provide cooling of said combustor liner, said annular passage being upstream of a burner and in serial flow communication with said burner, and
  • a perforated structure connecting the combustor liner and the outer shell, said structure being perforated with a multiplicity of openings adapted to provide an outlet of air from said annular passage to said burner with a predetermined pressure drop.
  • the underlying idea of the present invention is to provide a flow equalizer in the form of said perforated structure connecting the combustor liner and the outer shell.
  • the purpose of the perforations is to provide a certain pressure drop to the air entering the burner from the cooling passage which serves to equalize the unevenness of air flow. This further provides possibilities for mixing between air and fuel, resulting in low emissions and low combustor pulsations.
  • the perforated structure connecting the hot combustor liner and the colder outer shell is adapted to be flexible to accommodate the thermal displacements between the two components, particularly in a radial direction, and hence obviates the need for complicated arrangements involving moving parts that were previously used.
  • said structure is a conical structure connecting the combustor liner with the outer shell, and perforated with said multiplicity of openings.
  • said structure is integral to the combustor liner and the outer shell.
  • the above embodiment is advantageous in being relatively less complicated to implement .
  • the perforated d structure is made of a thinner material than the outer shell. This provides added flexibility to said structure, to accommodate the thermal displacement between the combustor liner and the outer shell.
  • FIG 1 is a longitudinal sectional view of a gas turbine having a serially cooled combustor according to one embodiment of the present invention
  • FIG 2 is an enlarged perspective view of the illustrated flow equalizer .
  • the gas turbine 10 includes a compressor section 14, a combustor 12 and a turbine section 16.
  • the compressor section 14 provides compressed air to the combustor 14.
  • the flow-path of the compressed air is represented by a dotted arrow 26.
  • the combustor 14 includes a burner 36 which includes means 30 for introducing a fuel into the compressed air for combustion thereof, producing a flame 40.
  • the hot combustion gases from the flame 40 are utilized by the turbine section 16 to produce mechanical work.
  • the combustor 12 is a can type combustor comprising a plurality of combustion chambers or cans 18 (only one can shown) arranged circumferentially around a central axis 19.
  • the combustor 12 may include an annular combustion chamber positioned around the central axis 19, without departing from the essence of the invention.
  • each combustor can 18 comprises a combustor liner 20, which is a cylindrical shell around a can axis 24.
  • the combustor liner is in direct contact with the combustion gases from the flame 40, and hence needs to be cooled in order to be maintained at an operating temperature that meets its durability requirements.
  • the combustor liner 20 is serially cooled by the air from the compressor section 14, which is first led to a narrow cooling passage 25 around the combustor liner 20 before flowing into the burner 36.
  • the cooling passage 25 is an annular passage defined by the combustor liner 20 and an outer shell 22 around the combustor liner 20, and forms a portion of the flow-path 26 for air before it flows into the burner 36.
  • the outer shell 22 may also be referred to as a cooling shield.
  • serial flow communication between the cooling passage 25 and the burner 36 is provided by a perforated thermal flexible structure 32 connecting the combustor liner 20 and the outer shell 22 that provides an outlet of the air from the cooling passage 25 to the hood of the burner 36.
  • the thermal flexible structure 32 is perforated with several small openings 33 and serves as a flow equalizer which provides a required pressure drop to the air at the burner inlet to equalize the unevenness of air flow at the burner inlet.
  • the thickness of the structure 32, the diameter of the openings 33, and the spacing between consecutive openings may be calculated to achieve a predetermined pressure drop at the burner inlet, so as to equalize uneven air flow, both locally in the combustor can 18, and globally between combustor cans 18. This further provides possibilities for mixing between air and fuel, resulting in low emissions and low combustor pulsations.
  • the operating temperature of the outer shell 22 is generally much lower than that of the combustor liner 20.
  • the temperature of the combustor liner 20 may be as high as 800 0 C, while that of the outer shell 22 is generally around 500 0 C.
  • the thermal mismatch between the combustor liner 20 and the outer shell 22 results in unequal thermal expansion between these components, leading to a thermal displacement, particularly in the radial direction.
  • the thermal flexible structure 32 connecting the combustor liner 20 and the outer shell 22 is suitably adapted to withstand the high thermal stress arising out of the unequal thermal expansions, and hence accommodate the thermal displacement between the two components without the need for complicated moving parts for the same. Flexibility of the structure 32 may be enhanced if the structure 32 is made of a thinner material than the outer shell 22.
  • the thermal flexible structure 32 is a flexible cone connecting the outer shell 22 and the combustor liner 20. Such a structure for a flow equalizer is particularly advantageous when the air flow at the burner inlet has a radial swirl.
  • the thermal flexible structure 32 may be made of, for example, stainless steel. Further, in order to minimize structural complexity of the combustor 12, the thermal flexible structure 32 may be made integral to the combustor liner 20 and the outer shell 22.
  • the present invention provides a combustor for a gas turbine.
  • the combustor comprises an annular passage defined by a combustor liner and an outer shell.
  • the annular passage forms a first portion of a flow-path for air, to provide cooling of said combustor liner.
  • the annular passage is upstream of a burner and in serial flow communication with said burner.
  • the proposed combustor further includes a structure connecting the combustor liner and the outer shell, the structure being perforated with a multiplicity of openings adapted to provide an outlet of air from said annular passage to said burner with a predetermined pressure drop .

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)

Abstract

The present invention provides a combustor (12) for a gas turbine (10). The combustor (12) comprises an annular passage (25) defined by a combustor liner (20) and an outer shell (22). The annular passage (25) forms a first portion of a flow-path (26) for air, to provide cooling of said combustor liner (20). The annular passage (25) is upstream of a burner (36) and in serial flow communication with said burner (36). The proposed combustor (12) further includes a perforated structure (32) connecting the combustor liner (20) and the outer shell (22). The structure (32) is perforated with a multiplicity of openings (33) adapted to provide an outlet of air from said annular passage (25) to said burner (36) with a predetermined pressure drop.

Description

Description
Combustor for a gas turbine
The present invention relates to gas turbines, and in particular, to a serially cooled combustor for a gas turbine.
A gas turbine engine includes a compressor section, a combustor and a turbine section. The compressor section delivers compressed air to the combustor which is used for combustion in a burner as well as for cooling the liner wall of the combustor, which must be maintained at an operating temperature meeting a durability requirement. In a serially cooled combustor, air from the compressor is first led through a cooling passage for cooling the liner wall of the combustor before flowing to the burner inlet. This cooling passage is usually defined by the combustor liner and an outer shell, sometimes referred to as a cooling shield. However, such serial cooling leads to uneven inflow of air at the burner inlet causing combustor pulsations. This unevenness of air inflow to the burner is generally corrected by adjustment of the burner fuel profile to avoid such pulsations which results in unnecessarily high emissions.
Furthermore, in operation, the combustor liner is in direct contact with the hot combustion gases and is hence subjected to a much higher temperature than the outer shell. This causes a thermal displacement between the hot combustor liner and the colder outer shell, as a result of which the connection of the hot combustor liner with the colder outer shell creates high thermal stresses in these components. Such thermal displacements have been hitherto addressed by providing complicated moving parts for connecting the combustor liner and the outer shell.
The object of the present invention is to provide an improved combustor for a gas turbine. The above object is achieved by a combustor for a gas turbine, comprising:
- an annular passage defined by a combustor liner and an outer shell, said annular passage forming a first portion of a flow-path for air, to provide cooling of said combustor liner, said annular passage being upstream of a burner and in serial flow communication with said burner, and
- a perforated structure connecting the combustor liner and the outer shell, said structure being perforated with a multiplicity of openings adapted to provide an outlet of air from said annular passage to said burner with a predetermined pressure drop.
The underlying idea of the present invention is to provide a flow equalizer in the form of said perforated structure connecting the combustor liner and the outer shell. The purpose of the perforations is to provide a certain pressure drop to the air entering the burner from the cooling passage which serves to equalize the unevenness of air flow. This further provides possibilities for mixing between air and fuel, resulting in low emissions and low combustor pulsations. Further, the perforated structure connecting the hot combustor liner and the colder outer shell is adapted to be flexible to accommodate the thermal displacements between the two components, particularly in a radial direction, and hence obviates the need for complicated arrangements involving moving parts that were previously used.
In a preferred embodiment, said structure is a conical structure connecting the combustor liner with the outer shell, and perforated with said multiplicity of openings. Such an arrangement is particularly advantageous when the air flow at the burner inlet has a radial swirl.
In one embodiment, said structure is integral to the combustor liner and the outer shell. The above embodiment is advantageous in being relatively less complicated to implement . In a preferred embodiment, the perforated d structure is made of a thinner material than the outer shell. This provides added flexibility to said structure, to accommodate the thermal displacement between the combustor liner and the outer shell.
The present invention is further described hereinafter with reference to illustrated embodiments shown in the accompanying drawings, in which:
FIG 1 is a longitudinal sectional view of a gas turbine having a serially cooled combustor according to one embodiment of the present invention, and
FIG 2 is an enlarged perspective view of the illustrated flow equalizer .
Referring to FIG 1, a portion of a gas turbine 10 is illustrated in accordance with one embodiment of the present invention. The gas turbine 10 includes a compressor section 14, a combustor 12 and a turbine section 16. The compressor section 14 provides compressed air to the combustor 14. The flow-path of the compressed air is represented by a dotted arrow 26. The combustor 14 includes a burner 36 which includes means 30 for introducing a fuel into the compressed air for combustion thereof, producing a flame 40. The hot combustion gases from the flame 40 are utilized by the turbine section 16 to produce mechanical work. In the illustrated embodiment, the combustor 12 is a can type combustor comprising a plurality of combustion chambers or cans 18 (only one can shown) arranged circumferentially around a central axis 19. In an alternate embodiment, the combustor 12 may include an annular combustion chamber positioned around the central axis 19, without departing from the essence of the invention. As shown, each combustor can 18 comprises a combustor liner 20, which is a cylindrical shell around a can axis 24. The combustor liner is in direct contact with the combustion gases from the flame 40, and hence needs to be cooled in order to be maintained at an operating temperature that meets its durability requirements. In the illustrated embodiment, the combustor liner 20 is serially cooled by the air from the compressor section 14, which is first led to a narrow cooling passage 25 around the combustor liner 20 before flowing into the burner 36. The cooling passage 25 is an annular passage defined by the combustor liner 20 and an outer shell 22 around the combustor liner 20, and forms a portion of the flow-path 26 for air before it flows into the burner 36. The outer shell 22 may also be referred to as a cooling shield.
In accordance with the present invention, serial flow communication between the cooling passage 25 and the burner 36 is provided by a perforated thermal flexible structure 32 connecting the combustor liner 20 and the outer shell 22 that provides an outlet of the air from the cooling passage 25 to the hood of the burner 36. As shown in greater detail in FIG 2, the thermal flexible structure 32 is perforated with several small openings 33 and serves as a flow equalizer which provides a required pressure drop to the air at the burner inlet to equalize the unevenness of air flow at the burner inlet. To that end, the thickness of the structure 32, the diameter of the openings 33, and the spacing between consecutive openings may be calculated to achieve a predetermined pressure drop at the burner inlet, so as to equalize uneven air flow, both locally in the combustor can 18, and globally between combustor cans 18. This further provides possibilities for mixing between air and fuel, resulting in low emissions and low combustor pulsations.
Further, it should be noted that the operating temperature of the outer shell 22 is generally much lower than that of the combustor liner 20. For example, in operation, the temperature of the combustor liner 20 may be as high as 8000C, while that of the outer shell 22 is generally around 5000C. The thermal mismatch between the combustor liner 20 and the outer shell 22 results in unequal thermal expansion between these components, leading to a thermal displacement, particularly in the radial direction. Accordingly, the thermal flexible structure 32 connecting the combustor liner 20 and the outer shell 22 is suitably adapted to withstand the high thermal stress arising out of the unequal thermal expansions, and hence accommodate the thermal displacement between the two components without the need for complicated moving parts for the same. Flexibility of the structure 32 may be enhanced if the structure 32 is made of a thinner material than the outer shell 22.
In the preferred embodiment illustrated in FIG 2, the thermal flexible structure 32 is a flexible cone connecting the outer shell 22 and the combustor liner 20. Such a structure for a flow equalizer is particularly advantageous when the air flow at the burner inlet has a radial swirl. The thermal flexible structure 32 may be made of, for example, stainless steel. Further, in order to minimize structural complexity of the combustor 12, the thermal flexible structure 32 may be made integral to the combustor liner 20 and the outer shell 22.
Summarizing, the present invention provides a combustor for a gas turbine. The combustor comprises an annular passage defined by a combustor liner and an outer shell. The annular passage forms a first portion of a flow-path for air, to provide cooling of said combustor liner. The annular passage is upstream of a burner and in serial flow communication with said burner. The proposed combustor further includes a structure connecting the combustor liner and the outer shell, the structure being perforated with a multiplicity of openings adapted to provide an outlet of air from said annular passage to said burner with a predetermined pressure drop . Although the invention has been described with reference to specific embodiments, this description is not meant to be construed in a limiting sense. Various modifications of the disclosed embodiments, as well as alternate embodiments of the invention, will become apparent to persons skilled in the art upon reference to the description of the invention. It is therefore contemplated that such modifications can be made without departing from the spirit or scope of the present invention as defined.

Claims

Patent claims
1. A combustor (12) for a gas turbine (10), comprising:
- an annular passage (25) defined by a combustor liner (20) and an outer shell (22), said annular passage (25) forming a first portion of a flow-path (26) for air, to provide cooling of said combustor liner (20), said annular passage (25) being upstream of a burner (36) and in serial flow communication with said burner (36), and
- a perforated structure (32) connecting the combustor liner (20) and the outer shell (22), said structure (32) being perforated with a multiplicity of openings (33) adapted to provide an outlet of air from said annular passage (25) to said burner
(36) with a predetermined pressure drop.
2. The combustor according to claim 1, wherein said structure (32) is a conical structure connecting the combustor liner (20) with the outer shell (22) and perforated with said multiplicity of openings (33) .
3. The combustor according to any of the preceding claims, wherein said structure (32) is integral to the combustor liner (20) and the outer shell (22) .
4. The combustor according to any of the preceding claims, wherein said structure (32) is made of a thinner material than the outer shell (22) .
PCT/EP2009/053701 2008-04-01 2009-03-30 Combustor for a gas turbine WO2009121820A1 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
EP08006656.6 2008-04-01
EP20080006656 EP2107314A1 (en) 2008-04-01 2008-04-01 Combustor for a gas turbine

Publications (1)

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WO2009121820A1 true WO2009121820A1 (en) 2009-10-08

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WO (1) WO2009121820A1 (en)

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
RU2667849C2 (en) * 2013-10-04 2018-09-24 Снекма Turbomachine combustion chamber provided with air deflection means for reducing wake created by ignition plug
US11156164B2 (en) 2019-05-21 2021-10-26 General Electric Company System and method for high frequency accoustic dampers with caps
US11174792B2 (en) 2019-05-21 2021-11-16 General Electric Company System and method for high frequency acoustic dampers with baffles

Families Citing this family (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JP6246562B2 (en) * 2013-11-05 2017-12-13 三菱日立パワーシステムズ株式会社 Gas turbine combustor

Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6220034B1 (en) * 1993-07-07 2001-04-24 R. Jan Mowill Convectively cooled, single stage, fully premixed controllable fuel/air combustor

Patent Citations (1)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US6220034B1 (en) * 1993-07-07 2001-04-24 R. Jan Mowill Convectively cooled, single stage, fully premixed controllable fuel/air combustor

Cited By (3)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
RU2667849C2 (en) * 2013-10-04 2018-09-24 Снекма Turbomachine combustion chamber provided with air deflection means for reducing wake created by ignition plug
US11156164B2 (en) 2019-05-21 2021-10-26 General Electric Company System and method for high frequency accoustic dampers with caps
US11174792B2 (en) 2019-05-21 2021-11-16 General Electric Company System and method for high frequency acoustic dampers with baffles

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Publication number Publication date
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