US20070089419A1 - Combustor for gas turbine engine - Google Patents

Combustor for gas turbine engine Download PDF

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Publication number
US20070089419A1
US20070089419A1 US11/582,295 US58229506A US2007089419A1 US 20070089419 A1 US20070089419 A1 US 20070089419A1 US 58229506 A US58229506 A US 58229506A US 2007089419 A1 US2007089419 A1 US 2007089419A1
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Prior art keywords
fuel supply
additive
premixed
combustor
fuel
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US11/582,295
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Kiyoshi Matsumoto
Atsushi Okuto
Yasuhiro Kinoshita
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Kawasaki Motors Ltd
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Kawasaki Jukogyo KK
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Assigned to KAWASAKI JUKOGYO KABUSHIKI KAISHA reassignment KAWASAKI JUKOGYO KABUSHIKI KAISHA ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: KINOSHITA, YASUHIRO, MATSUMOTO, KIYOSHI, OKUTO, ATSUSHI
Publication of US20070089419A1 publication Critical patent/US20070089419A1/en
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    • 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
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/34—Feeding into different combustion zones
    • F23R3/346—Feeding into different combustion zones for staged combustion
    • 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
    • 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
    • F23R3/10—Air inlet arrangements for primary air
    • F23R3/12—Air inlet arrangements for primary air inducing a vortex
    • F23R3/14—Air inlet arrangements for primary air inducing a vortex by using swirl vanes

Definitions

  • the present invention relates to a combustor for a gas turbine engine. More specifically, the present invention relates to a combustor for a gas turbine engine, capable of making the combustor of the gas turbine engine achieve desired stable, low-pollution combustion under partial load.
  • Private power generation equipment has progressively come into wide use recently, and gas turbine engines are used widely as power units for private power generation equipment. Different time zones have different power demands, respectively, on private power generation equipment. Therefore, private power generation equipment does not operate always in a maximum-load operating mode and are often required to operate in a partial-load operating mode.
  • the gas turbine engine for private power generation equipment is required to maintain satisfactorily stable combustion and to reduce the rate of emission of harmful gases including NO x (nitrogen oxides) and CO (carbon monoxide) for environmental conservation.
  • NO x nitrogen oxides
  • CO carbon monoxide
  • the first method needs an additional mechanism for injecting water or steam into the combustion chamber.
  • the additional mechanism enlarges the gas turbine engine and the gas turbine engine cannot reduce the rate of NO x emission greatly so as to compensate for the disadvantage of the additional mechanism.
  • the second method can reduce the rate of NO x emission in a wide range of load on the gas turbine engine even in lean premixed combustion having a narrow stable combustion range, work necessary for extraction reduces the thermal efficiency of the gas turbine engine.
  • the present invention has been made in view of those problems in the related art and it is therefore an object of the present invention to provide a combustor for a gas turbine engine, capable of achieving stable combustion and of reducing the rate of emission of harmful substances, such as NO x and CO, in a wide range of load without enlarging the gas turbine engine, without complicating the constitution of the gas turbine engine and without reducing the thermal efficiency of the gas turbine engine.
  • a combustor for a gas turbine engine of a lean premixed combustion type includes: a fuel jetting device configured to jet a premixed fuel into a combustion chamber so that the premixed fuel burns in a premixed combustion zone which is formed in the combustion chamber; and an additive fuel supply unit disposed downstream of the premixed combustion zone so as to additionally supply fuel into the combustion chamber, the additive fuel supply unit including first and second additive fuel supply devices disposed in a two-stage arrangement such that the second additive fuel supply device is positioned downstream of the first additive fuel supply device.
  • a main air passage formed in a liner head to supply air into the premixed combustion zone has a sectional area not lager than a sum of sectional areas of air passages in the first and second additive fuel supply devices of the additive fuel supply unit.
  • a sectional area of an air passage in the first additive fuel supply device is not larger than a sectional area of an air passage in the second additive fuel supply device.
  • the first and second additive fuel supply devices of the additive fuel supply unit are respectively provided with diffusion nozzles to supply fuel.
  • the combustor according to the present invention for a gas turbine engine achieves desired stable, low-pollution combustion under load varying in a wide load range and discharges an exhaust gas including reduced amount of harmful substances, such as NO x and CO.
  • FIG. 1 is a schematic sectional view of a combustor U in a preferred embodiment according to the present invention for a gas turbine engine
  • FIG. 2 is diagram comparatively showing the respective load characteristics of harmful gas emissions of a combustor in a present example of the present invention and of a combustor in a comparative example.
  • a combustor in a preferred embodiment according to the present invention will be described by way of example with reference to FIGS. 1 and 2 .
  • the combustor U has a combustor casing including a flow sleeve 1 , namely, an outer sleeve, and an inner tube 2 , namely, an inner sleeve, extended inside the flow sleeve 1 .
  • One end of the flow sleeve 1 and one end of the inner tube 2 are closed by an end wall 3 .
  • the other end of the flow sleeve 1 is an air inlet 1 a and the other end of the inner tube 2 is an exhaust outlet 2 a.
  • the interior of the inner tube 2 is divided into a liner head 4 on the side of the end wall 3 and a combustion chamber 5 on the side of the exhaust outlet 2 a by a partition wall 2 b having a shape resembling an annular plate disposed at a predetermined longitudinal position.
  • the liner head 4 is provided with a predetermined number of openings 4 a through which combustion air taken through the air inlet 1 a into the flow sleeve 1 flows into the inner tube 2 .
  • a fuel jetting device 6 is disposed inside the liner head 4 .
  • the fuel jetting device 6 prepares a premixed fuel by mixing fuel supplied by a fuel supply system, not shown, and combustion air and jets the premixed fuel into the combustion chamber 5 .
  • the premixed fuel jetted by the fuel jetting device 6 burns in a predetermined premixed combustion zone V in the combustion chamber 5 .
  • the combustor U is provided with a two-stage additive fuel supply unit 7 in addition to the fuel jetting device 6 .
  • the two-stage additive fuel supply unit 7 supplies fuel additionally into the combustion chamber 5 while the gas turbine engine is in a partial-load operation.
  • a character F indicates diffusion flames, namely, premixed flames.
  • the fuel jetting device 6 includes a pilot combustion unit 10 in a central part of an inner opening of the partition wall 2 b disposed on the upstream side of the combustion chamber 5 , and a main combustion unit 20 surrounding the pilot combustion unit 10 .
  • the pilot combustion unit 10 has a central pilot fuel nozzle 11 , and a combustion air passage 12 surrounding the pilot fuel nozzle 11 .
  • a premixed fuel produced by mixing fuel jetted through the pilot fuel nozzle 11 and combustion air is jetted into the combustion chamber 5 .
  • the pilot fuel nozzle 11 has a closed end and a side wall provided with radial jetting holes 11 a. Fuel is jetted radially through the radial jetting holes 11 a.
  • the combustion air passage 12 is defined by the outside surface of the pilot fuel nozzle 11 and the inside surface of an inner hat-shaped tube 13 surrounding the pilot fuel nozzle 11 .
  • a swirler 14 is disposed at a predetermined position in the combustion air passage 12 .
  • the fuel jetted into the combustion air passage 12 through the pilot fuel nozzle 11 is mixed into swirling combustion air currents produced by the swirler 14 .
  • the premixed fuel flows into the combustion chamber 5 .
  • the main combustion unit 20 includes a main fuel nozzle 21 and a main premixed fuel passage 22 .
  • the main fuel nozzle 21 surrounds the inner hat-shaped tube 13 .
  • the fuel is jetted through the main fuel nozzle 21 toward the inlet 22 a of the premixed fuel passage 22 .
  • the main premixed fuel passage 22 is defined by the outside surface of the inner hat-shaped tube 13 and an outer hat-shaped tube 23 of a diameter greater than that of the inner hat-shaped tube 13 .
  • a space between the flanges of the hat-shaped tubes 13 and 23 forms the inlet 22 a of the premixed fuel passage 22 .
  • a swirler 24 is disposed in the inlet 22 a of the premixed fuel passage 22 .
  • the open end of the outer hat-shaped tube 23 is joined to the inside edge of the partition wall 2 b disposed inside the inner tube 2 to form a premixed fuel outlet 22 b , namely, the outlet of the main premixed fuel passage.
  • a premixed fuel produced by mixing the fuel jetted through the main fuel nozzle 21 and combustion air is swirled.
  • the swirling premixed fuel is jetted through the premixed fuel outlet 22 b into the premixed combustion zone V.
  • the present specification defines the sum of the respective sectional areas of the main premixed fuel passage 22 and the swirler 14 as a main air passage area S.
  • the main air passage area S is determined such that the air-fuel ratio of the premixed fuel in the premixed combustion zone V while the gas turbine engine is in a partial-load operation under a predetermined load, such as a load equal to 50% of a maximum load, is nearly equal to the limited air-fuel ratio for an operation under a predetermined load nearly equal to a maximum load, such as a load equal to 80% of the maximum load.
  • the main air passage area S is smaller than an ordinary main air passage area.
  • the two-stage additive fuel supply unit 7 supplies fuel additionally for after burning into a predetermined zone downstream of the premixed combustion zone V, in which stable low-NO x , low-CO combustion is carried out, in the combustion chamber 5 during a partial-load operation to suppress the generation of NO x .
  • the two-stage additive fuel supply unit 7 has a predetermined number of additive fuel supply tubes 31 penetrate parts corresponding to the predetermined zone of the inner tube 2 defining the combustion chamber 5 , and additive fuel supply nozzles 32 having end parts of a predetermined length inserted into the additive fuel supply tubes 31 .
  • the additive fuel supply nozzles 32 may be either of diffusion nozzles as shown in FIG. 1 and premixed combustion nozzles, not shown.
  • the two-stage additive fuel supply unit 7 includes a first additive fuel supply device 7 A and a second additive fuel supply device 7 B.
  • the first additive fuel supply device 7 a and the second additive fuel supply device 7 B are disposed at a first longitudinal distance L 1 and a second longitudinal distance L 2 , respectively, from the premixed fuel outlet 22 b .
  • the distance L 2 is greater than the distance L 1 .
  • the first distance L 1 , the second distance L 2 , the numbers of the tubes 31 and the additive fuel supply nozzles 32 of the additive fuel supply devices 7 a and 7 B, the pattern of arrangement, such as a straight pattern or a zigzag pattern, of the additive fuel supply nozzles 32 , the ratio between the main air passage area S and the sum of the sectional areas of the air passages of the two-stage additive fuel supply unit 7 , and the ratio between the respective sectional areas of the first additive fuel supply device 7 A and the second additive fuel supply device 7 B are selectively determined so that the rate of emission of NO x and CO may be small and stable combustion can be achieved.
  • the ratio D 1 of the main air passage area S to the total sectional area of the two-stage additive fuel supply unit 7 is, for example, 5/5
  • the ratio D 2 of the total sectional area of the first additive fuel supply device 7 A to the total sectional area of the second additive fuel supply device 7 B is, for example, 4/6.
  • the ratio D 1 is reduced to 4/6 or such to widen the range of low-load operation.
  • the ratio D 2 is reduced to 3/7 or such to widen the range of low-load operation.
  • the main air passage area S of the combustor U in this embodiment is determined so that the combustor U can achieve stable combustion at a low rate of emission of NO x and CO while the gas turbine engine is in a partial-load operation, and the fuel additionally supplied by the two-stage additive fuel supply unit 7 burns in the combustion chamber 5 . Therefore, a proper air-fuel ratio in accordance with a given load can be determined so that the generation of NO x and CO can be suppressed.
  • the combustor U can maintain stable combustion at a low rate of emission of NO x and CO for the operation of the gas turbine in a wide load range.
  • FIG. 2 Shown in FIG. 2 are load characteristics of NO x and CO concentrations of the exhaust gas discharged from a combustor in the comparative example provided with only the first additive fuel supply device 7 A and load characteristics of NO x and CO concentrations of the exhaust gas discharged from a combustor in the present example provided with the first additive fuel supply device 7 A and the second additive fuel supply device 7 B.
  • H 1 shows the allowable upper limit of NO x concentration
  • H 2 shows the allowable upper limit of CO concentration.
  • the NO x concentration of the exhaust gas discharged from the combustor in the present example is lower than that of the exhaust gas discharged from the combustor in the comparative example in the entire load range.
  • the additive fuel supply devices 7 A and 7 B are effective in reducing the NO x and the CO concentration of the exhaust gas.
  • the CO concentration of the exhaust gas discharged from the combustor in the present example is below the allowable upper limit H 2 of CO concentration in the entire load range, which proves that the additive fuel supply devices 7 A and 7 B are effective in reducing the CO concentration of the exhaust gas.
  • the CO concentration of the exhaust gas generated by the combustor in the comparative example is higher than the allowable upper limit H 2 of CO concentration at loads below a predetermined load E 1 . Therefore, the combustor in the comparative example should not be operated at loads below the predetermined load E 1 from the viewpoint of environmental conservation.
  • the two-stage additive fuel supply unit 7 including the additive fuel supply devices 7 A and 7 B expands the load range in which the combustor can operate without causing problems in environmental conservation.

Abstract

A combustor for a gas turbine engine of a lean premixed combustion type, includes: a fuel jetting device configured to jet a premixed fuel into a combustion chamber so that the premixed fuel burns in a premixed combustion zone which is formed in the combustion chamber; and an additive fuel supply unit disposed downstream of the premixed combustion zone so as to additionally supply fuel into the combustion chamber. The additive fuel supply unit includes first and second additive fuel supply devices disposed in a two-stage arrangement such that the second additive fuel supply device is positioned downstream of the first additive fuel supply device.

Description

    CROSS REFERENCE TO RELATED APPLICATIONS
  • This application is based upon the prior Japanese Patent Application No. 2005-308380 filed on Oct. 24, 2005, the entire contents of which are incorporated herein by reference.
    • BACKGROUND OF THE INVENTION
  • 1. Field of the Invention
  • The present invention relates to a combustor for a gas turbine engine. More specifically, the present invention relates to a combustor for a gas turbine engine, capable of making the combustor of the gas turbine engine achieve desired stable, low-pollution combustion under partial load.
  • 2. Description of the Related Art
  • Private power generation equipment has progressively come into wide use recently, and gas turbine engines are used widely as power units for private power generation equipment. Different time zones have different power demands, respectively, on private power generation equipment. Therefore, private power generation equipment does not operate always in a maximum-load operating mode and are often required to operate in a partial-load operating mode.
  • Consequently, the gas turbine engine for private power generation equipment is required to maintain satisfactorily stable combustion and to reduce the rate of emission of harmful gases including NOx (nitrogen oxides) and CO (carbon monoxide) for environmental conservation.
  • In order to meet the requirements, some techniques have been proposed previously, such as a first method that injects water or steam into a combustion chamber and a second method that reduces air for lean premixed combustion through engine extraction.
  • The first method needs an additional mechanism for injecting water or steam into the combustion chamber. The additional mechanism enlarges the gas turbine engine and the gas turbine engine cannot reduce the rate of NOx emission greatly so as to compensate for the disadvantage of the additional mechanism.
  • Although the second method can reduce the rate of NOx emission in a wide range of load on the gas turbine engine even in lean premixed combustion having a narrow stable combustion range, work necessary for extraction reduces the thermal efficiency of the gas turbine engine.
  • The previously proposed techniques are described in JP-A 8-261468, JP-A 10-196909 and JP-A 2000-274689.
    • SUMMARY OF THE INVENTION
  • The present invention has been made in view of those problems in the related art and it is therefore an object of the present invention to provide a combustor for a gas turbine engine, capable of achieving stable combustion and of reducing the rate of emission of harmful substances, such as NOx and CO, in a wide range of load without enlarging the gas turbine engine, without complicating the constitution of the gas turbine engine and without reducing the thermal efficiency of the gas turbine engine.
  • A combustor for a gas turbine engine of a lean premixed combustion type according to the present invention includes: a fuel jetting device configured to jet a premixed fuel into a combustion chamber so that the premixed fuel burns in a premixed combustion zone which is formed in the combustion chamber; and an additive fuel supply unit disposed downstream of the premixed combustion zone so as to additionally supply fuel into the combustion chamber, the additive fuel supply unit including first and second additive fuel supply devices disposed in a two-stage arrangement such that the second additive fuel supply device is positioned downstream of the first additive fuel supply device.
  • In the combustor according to the present invention for a gas turbine engine, it is preferable that a main air passage formed in a liner head to supply air into the premixed combustion zone has a sectional area not lager than a sum of sectional areas of air passages in the first and second additive fuel supply devices of the additive fuel supply unit.
  • In the combustor according to the present invention for a gas turbine engine, it is preferable that a sectional area of an air passage in the first additive fuel supply device is not larger than a sectional area of an air passage in the second additive fuel supply device.
  • In the combustor according to the present invention for a gas turbine engine, it is preferable that the first and second additive fuel supply devices of the additive fuel supply unit are respectively provided with diffusion nozzles to supply fuel.
  • The combustor according to the present invention for a gas turbine engine achieves desired stable, low-pollution combustion under load varying in a wide load range and discharges an exhaust gas including reduced amount of harmful substances, such as NOx and CO.
    • BRIEF DESCRIPTION OF THE DRAWINGS
  • The above and other objects, features and advantages of the present invention will become more apparent from the following description taken in connection with the accompanying drawings, in which:
  • FIG. 1 is a schematic sectional view of a combustor U in a preferred embodiment according to the present invention for a gas turbine engine; and
  • FIG. 2 is diagram comparatively showing the respective load characteristics of harmful gas emissions of a combustor in a present example of the present invention and of a combustor in a comparative example.
    • DESCRIPTION OF THE PREFERRED EMBODIMENTS
  • A combustor in a preferred embodiment according to the present invention will be described by way of example with reference to FIGS. 1 and 2.
  • Referring to FIG. 1 showing a combustor U in a preferred embodiment according to the present invention included in a gas turbine engine, the combustor U has a combustor casing including a flow sleeve 1, namely, an outer sleeve, and an inner tube 2, namely, an inner sleeve, extended inside the flow sleeve 1.
  • One end of the flow sleeve 1 and one end of the inner tube 2 are closed by an end wall 3. The other end of the flow sleeve 1 is an air inlet 1 a and the other end of the inner tube 2 is an exhaust outlet 2 a.
  • The interior of the inner tube 2 is divided into a liner head 4 on the side of the end wall 3 and a combustion chamber 5 on the side of the exhaust outlet 2 a by a partition wall 2 b having a shape resembling an annular plate disposed at a predetermined longitudinal position.
  • The liner head 4 is provided with a predetermined number of openings 4 a through which combustion air taken through the air inlet 1 a into the flow sleeve 1 flows into the inner tube 2. A fuel jetting device 6 is disposed inside the liner head 4. The fuel jetting device 6 prepares a premixed fuel by mixing fuel supplied by a fuel supply system, not shown, and combustion air and jets the premixed fuel into the combustion chamber 5. The premixed fuel jetted by the fuel jetting device 6 burns in a predetermined premixed combustion zone V in the combustion chamber 5.
  • The combustor U is provided with a two-stage additive fuel supply unit 7 in addition to the fuel jetting device 6. The two-stage additive fuel supply unit 7 supplies fuel additionally into the combustion chamber 5 while the gas turbine engine is in a partial-load operation. In FIG. 1, a character F indicates diffusion flames, namely, premixed flames.
  • The fuel jetting device 6 includes a pilot combustion unit 10 in a central part of an inner opening of the partition wall 2 b disposed on the upstream side of the combustion chamber 5, and a main combustion unit 20 surrounding the pilot combustion unit 10.
  • The pilot combustion unit 10 has a central pilot fuel nozzle 11, and a combustion air passage 12 surrounding the pilot fuel nozzle 11. A premixed fuel produced by mixing fuel jetted through the pilot fuel nozzle 11 and combustion air is jetted into the combustion chamber 5.
  • The pilot fuel nozzle 11 has a closed end and a side wall provided with radial jetting holes 11 a. Fuel is jetted radially through the radial jetting holes 11 a. The combustion air passage 12 is defined by the outside surface of the pilot fuel nozzle 11 and the inside surface of an inner hat-shaped tube 13 surrounding the pilot fuel nozzle 11. A swirler 14 is disposed at a predetermined position in the combustion air passage 12. The fuel jetted into the combustion air passage 12 through the pilot fuel nozzle 11 is mixed into swirling combustion air currents produced by the swirler 14. The premixed fuel flows into the combustion chamber 5.
  • The main combustion unit 20 includes a main fuel nozzle 21 and a main premixed fuel passage 22. A premixed fuel produced by mixing the fuel jetted through the main fuel nozzle 21 and combustion air flows through the main premixed fuel passage 22 into the combustion chamber 5.
  • The main fuel nozzle 21 surrounds the inner hat-shaped tube 13. The fuel is jetted through the main fuel nozzle 21 toward the inlet 22 a of the premixed fuel passage 22.
  • The main premixed fuel passage 22 is defined by the outside surface of the inner hat-shaped tube 13 and an outer hat-shaped tube 23 of a diameter greater than that of the inner hat-shaped tube 13. A space between the flanges of the hat- shaped tubes 13 and 23 forms the inlet 22 a of the premixed fuel passage 22.
  • A swirler 24 is disposed in the inlet 22 a of the premixed fuel passage 22.
  • The open end of the outer hat-shaped tube 23 is joined to the inside edge of the partition wall 2 b disposed inside the inner tube 2 to form a premixed fuel outlet 22 b, namely, the outlet of the main premixed fuel passage.
  • A premixed fuel produced by mixing the fuel jetted through the main fuel nozzle 21 and combustion air is swirled. The swirling premixed fuel is jetted through the premixed fuel outlet 22 b into the premixed combustion zone V.
  • Combustion air taken through the openings 4 a of the liner head 4 into the inner tube 2 flows through the main premixed fuel passage 22 and the swirler 14 into the premixed combustion zone V. The present specification defines the sum of the respective sectional areas of the main premixed fuel passage 22 and the swirler 14 as a main air passage area S.
  • Generally, there are limited lean air-fuel ratios that enable the gas turbine engine to achieve stable combustion at low rate of emission of NOx and CO under low loads. According to the present embodiment, the main air passage area S is determined such that the air-fuel ratio of the premixed fuel in the premixed combustion zone V while the gas turbine engine is in a partial-load operation under a predetermined load, such as a load equal to 50% of a maximum load, is nearly equal to the limited air-fuel ratio for an operation under a predetermined load nearly equal to a maximum load, such as a load equal to 80% of the maximum load. The main air passage area S is smaller than an ordinary main air passage area.
  • The two-stage additive fuel supply unit 7 supplies fuel additionally for after burning into a predetermined zone downstream of the premixed combustion zone V, in which stable low-NOx, low-CO combustion is carried out, in the combustion chamber 5 during a partial-load operation to suppress the generation of NOx. The two-stage additive fuel supply unit 7 has a predetermined number of additive fuel supply tubes 31 penetrate parts corresponding to the predetermined zone of the inner tube 2 defining the combustion chamber 5, and additive fuel supply nozzles 32 having end parts of a predetermined length inserted into the additive fuel supply tubes 31.
  • The additive fuel supply nozzles 32 may be either of diffusion nozzles as shown in FIG. 1 and premixed combustion nozzles, not shown.
  • The two-stage additive fuel supply unit 7 includes a first additive fuel supply device 7A and a second additive fuel supply device 7B. The first additive fuel supply device 7 a and the second additive fuel supply device 7B are disposed at a first longitudinal distance L1 and a second longitudinal distance L2, respectively, from the premixed fuel outlet 22 b. The distance L2 is greater than the distance L1.
  • The first distance L1, the second distance L2, the numbers of the tubes 31 and the additive fuel supply nozzles 32 of the additive fuel supply devices 7 a and 7B, the pattern of arrangement, such as a straight pattern or a zigzag pattern, of the additive fuel supply nozzles 32, the ratio between the main air passage area S and the sum of the sectional areas of the air passages of the two-stage additive fuel supply unit 7, and the ratio between the respective sectional areas of the first additive fuel supply device 7A and the second additive fuel supply device 7B are selectively determined so that the rate of emission of NOx and CO may be small and stable combustion can be achieved.
  • The ratio D1 of the main air passage area S to the total sectional area of the two-stage additive fuel supply unit 7 is, for example, 5/5, and the ratio D2 of the total sectional area of the first additive fuel supply device 7A to the total sectional area of the second additive fuel supply device 7B is, for example, 4/6.
  • The ratio D1 is reduced to 4/6 or such to widen the range of low-load operation.
  • The ratio D2 is reduced to 3/7 or such to widen the range of low-load operation.
  • The main air passage area S of the combustor U in this embodiment is determined so that the combustor U can achieve stable combustion at a low rate of emission of NOx and CO while the gas turbine engine is in a partial-load operation, and the fuel additionally supplied by the two-stage additive fuel supply unit 7 burns in the combustion chamber 5. Therefore, a proper air-fuel ratio in accordance with a given load can be determined so that the generation of NOx and CO can be suppressed.
  • Thus the combustor U can maintain stable combustion at a low rate of emission of NOx and CO for the operation of the gas turbine in a wide load range.
  • Performance of combustors in a present example according to the present invention and a comparative example will be described in connection with FIG. 2. Shown in FIG. 2 are load characteristics of NOx and CO concentrations of the exhaust gas discharged from a combustor in the comparative example provided with only the first additive fuel supply device 7A and load characteristics of NOx and CO concentrations of the exhaust gas discharged from a combustor in the present example provided with the first additive fuel supply device 7A and the second additive fuel supply device 7B. In FIG. 2, H1 shows the allowable upper limit of NOx concentration, and H2 shows the allowable upper limit of CO concentration.
  • As obvious from FIG. 2, the NOx concentration of the exhaust gas discharged from the combustor in the present example is lower than that of the exhaust gas discharged from the combustor in the comparative example in the entire load range.
  • It is known from FIG. 2 that the additive fuel supply devices 7A and 7B are effective in reducing the NOx and the CO concentration of the exhaust gas. The CO concentration of the exhaust gas discharged from the combustor in the present example is below the allowable upper limit H2 of CO concentration in the entire load range, which proves that the additive fuel supply devices 7A and 7B are effective in reducing the CO concentration of the exhaust gas. The CO concentration of the exhaust gas generated by the combustor in the comparative example is higher than the allowable upper limit H2 of CO concentration at loads below a predetermined load E1. Therefore, the combustor in the comparative example should not be operated at loads below the predetermined load E1 from the viewpoint of environmental conservation. Thus the two-stage additive fuel supply unit 7 including the additive fuel supply devices 7A and 7B expands the load range in which the combustor can operate without causing problems in environmental conservation.
  • Although the invention has been described in its preferred embodiment with a certain degree of particularity, obviously many changes and variations are possible therein. It is therefore to be understood that the present invention may be practiced otherwise than as specifically described herein without departing from the scope and spirit thereof.

Claims (5)

1. A combustor for a gas turbine engine of a lean premixed combustion type, comprising:
a fuel jetting device configured to jet a premixed fuel into a combustion chamber so that the premixed fuel burns in a premixed combustion zone which is formed in the combustion chamber; and
an additive fuel supply unit disposed downstream of the premixed combustion zone so as to additionally supply fuel into the combustion chamber, the additive fuel supply unit including first and second additive fuel supply devices disposed in a two-stage arrangement such that the second additive fuel supply device is positioned downstream of the first additive fuel supply device.
2. The combustor for a gas turbine engine according to claim 1, wherein a main air passage formed in a liner head to supply air into the premixed combustion zone has a sectional area not lager than a sum of sectional areas of air passages in the first and second additive fuel supply devices of the additive fuel supply unit.
3. The combustor for a gas turbine engine according to claim 1, wherein a sectional area of an air passage in the first additive fuel supply device is not larger than a sectional area of an air passage in the second additive fuel supply device.
4. The combustor for a gas turbine engine according to claim 2, wherein a sectional area of an air passage in the first additive fuel supply device is not larger than a sectional area of an air passage in the second additive fuel supply device.
5. The combustor for a gas turbine engine according to claim 1, wherein the first and second additive fuel supply devices of the additive fuel supply unit are respectively provided with diffusion nozzles to supply fuel.
US11/582,295 2005-10-24 2006-10-18 Combustor for gas turbine engine Abandoned US20070089419A1 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
JP2005308380A JP2007113888A (en) 2005-10-24 2005-10-24 Combustor structure of gas turbine engine
JP2005-308380 2005-10-24

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US20150167980A1 (en) * 2013-12-18 2015-06-18 Jared M. Pent Axial stage injection dual frequency resonator for a combustor of a gas turbine engine
US20150184858A1 (en) * 2012-10-01 2015-07-02 Peter John Stuttford Method of operating a multi-stage flamesheet combustor
US20150275755A1 (en) * 2012-12-13 2015-10-01 Kawasaki Jukogyo Kabushiki Kaisha Multi-fuel-capable gas turbine combustor
CN105423344A (en) * 2016-01-12 2016-03-23 西北工业大学 Flame tube structure of combustion chamber of ground gas turbine
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US9400113B2 (en) 2014-06-12 2016-07-26 Kawasaki Jukogyo Kabushiki Kaisha Multifuel gas turbine combustor
US20160258629A1 (en) * 2015-03-06 2016-09-08 General Electric Company Fuel staging in a gas turbine engine
US9752781B2 (en) 2012-10-01 2017-09-05 Ansaldo Energia Ip Uk Limited Flamesheet combustor dome
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US20170307210A1 (en) * 2014-12-02 2017-10-26 Kawasaki Jukogyo Kabushiki Kaisha Gas turbine combustor and gas turbine
US9897317B2 (en) 2012-10-01 2018-02-20 Ansaldo Energia Ip Uk Limited Thermally free liner retention mechanism
US20180229927A1 (en) * 2015-08-07 2018-08-16 Red Valve Company, Inc. Free Flow Collar for Control Valves
US10060630B2 (en) 2012-10-01 2018-08-28 Ansaldo Energia Ip Uk Limited Flamesheet combustor contoured liner
US10378456B2 (en) 2012-10-01 2019-08-13 Ansaldo Energia Switzerland AG Method of operating a multi-stage flamesheet combustor
US10584880B2 (en) * 2016-03-25 2020-03-10 General Electric Company Mounting of integrated combustor nozzles in a segmented annular combustion system
US10584876B2 (en) * 2016-03-25 2020-03-10 General Electric Company Micro-channel cooling of integrated combustor nozzle of a segmented annular combustion system
US10605459B2 (en) * 2016-03-25 2020-03-31 General Electric Company Integrated combustor nozzle for a segmented annular combustion system
US11187415B2 (en) * 2017-12-11 2021-11-30 General Electric Company Fuel injection assemblies for axial fuel staging in gas turbine combustors
US11255545B1 (en) 2020-10-26 2022-02-22 General Electric Company Integrated combustion nozzle having a unified head end
US11371702B2 (en) 2020-08-31 2022-06-28 General Electric Company Impingement panel for a turbomachine
US11460191B2 (en) 2020-08-31 2022-10-04 General Electric Company Cooling insert for a turbomachine
US11614233B2 (en) 2020-08-31 2023-03-28 General Electric Company Impingement panel support structure and method of manufacture
US20230228425A1 (en) * 2022-01-18 2023-07-20 Qingdao Zhennuo Laser Technology Co., Ltd. Multi-Nozzle Fuel Injection Method for Gas Turbine
US11767766B1 (en) 2022-07-29 2023-09-26 General Electric Company Turbomachine airfoil having impingement cooling passages
US11846426B2 (en) * 2021-06-24 2023-12-19 General Electric Company Gas turbine combustor having secondary fuel nozzles with plural passages for injecting a diluent and a fuel
DE102022208339A1 (en) 2022-08-10 2024-02-15 Rolls-Royce Deutschland Ltd & Co Kg Combustion chamber arrangement for operation with liquid and/or gaseous fuel, gas turbine arrangement and method

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US8387398B2 (en) * 2007-09-14 2013-03-05 Siemens Energy, Inc. Apparatus and method for controlling the secondary injection of fuel
US20090084082A1 (en) * 2007-09-14 2009-04-02 Siemens Power Generation, Inc. Apparatus and Method for Controlling the Secondary Injection of Fuel
US20100242482A1 (en) * 2009-03-30 2010-09-30 General Electric Company Method and system for reducing the level of emissions generated by a system
US8689559B2 (en) * 2009-03-30 2014-04-08 General Electric Company Secondary combustion system for reducing the level of emissions generated by a turbomachine
US8596069B2 (en) * 2011-06-28 2013-12-03 General Electric Company Rational late lean injection
US20130180255A1 (en) * 2011-06-28 2013-07-18 General Electric Company Rational late lean injection
CN102954469A (en) * 2011-08-05 2013-03-06 通用电气公司 Assembly and apparatus related to integrating late lean injection into combustion turbine engine
US8919137B2 (en) * 2011-08-05 2014-12-30 General Electric Company Assemblies and apparatus related to integrating late lean injection into combustion turbine engines
US9388987B2 (en) * 2011-09-22 2016-07-12 General Electric Company Combustor and method for supplying fuel to a combustor
US20140338359A1 (en) * 2011-09-22 2014-11-20 General Electric Company Combustor and method for supplying fuel to a combustor
US20140238034A1 (en) * 2011-11-17 2014-08-28 General Electric Company Turbomachine combustor assembly and method of operating a turbomachine
US20130213046A1 (en) * 2012-02-16 2013-08-22 General Electric Company Late lean injection system
CN103256630A (en) * 2012-02-16 2013-08-21 通用电气公司 Late lean injection system
US9752781B2 (en) 2012-10-01 2017-09-05 Ansaldo Energia Ip Uk Limited Flamesheet combustor dome
US10060630B2 (en) 2012-10-01 2018-08-28 Ansaldo Energia Ip Uk Limited Flamesheet combustor contoured liner
US20150184858A1 (en) * 2012-10-01 2015-07-02 Peter John Stuttford Method of operating a multi-stage flamesheet combustor
US10378456B2 (en) 2012-10-01 2019-08-13 Ansaldo Energia Switzerland AG Method of operating a multi-stage flamesheet combustor
US9897317B2 (en) 2012-10-01 2018-02-20 Ansaldo Energia Ip Uk Limited Thermally free liner retention mechanism
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US20150275755A1 (en) * 2012-12-13 2015-10-01 Kawasaki Jukogyo Kabushiki Kaisha Multi-fuel-capable gas turbine combustor
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US20160123596A1 (en) * 2013-06-11 2016-05-05 United Technologies Corporation Combustor with axial staging for a gas turbine engine
US20150052905A1 (en) * 2013-08-20 2015-02-26 General Electric Company Pulse Width Modulation for Control of Late Lean Liquid Injection Velocity
US20150159877A1 (en) * 2013-12-06 2015-06-11 General Electric Company Late lean injection manifold mixing system
US20150167980A1 (en) * 2013-12-18 2015-06-18 Jared M. Pent Axial stage injection dual frequency resonator for a combustor of a gas turbine engine
US9638423B2 (en) * 2014-06-12 2017-05-02 Kawasaki Jukogyo Kabushiki Kaisha Multifuel gas turbine combustor with fuel mixing chamber and supplemental burner
US9400113B2 (en) 2014-06-12 2016-07-26 Kawasaki Jukogyo Kabushiki Kaisha Multifuel gas turbine combustor
US20170298817A1 (en) * 2014-10-10 2017-10-19 Kawasaki Jukogyo Kabushiki Kaisha Combustor and gas turbine engine
US20170307210A1 (en) * 2014-12-02 2017-10-26 Kawasaki Jukogyo Kabushiki Kaisha Gas turbine combustor and gas turbine
US20160258629A1 (en) * 2015-03-06 2016-09-08 General Electric Company Fuel staging in a gas turbine engine
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US10773878B2 (en) * 2015-08-07 2020-09-15 Red Valve Company, Inc. Free flow collar for control valves
US20180229927A1 (en) * 2015-08-07 2018-08-16 Red Valve Company, Inc. Free Flow Collar for Control Valves
CN105423344A (en) * 2016-01-12 2016-03-23 西北工业大学 Flame tube structure of combustion chamber of ground gas turbine
US10584880B2 (en) * 2016-03-25 2020-03-10 General Electric Company Mounting of integrated combustor nozzles in a segmented annular combustion system
US10584876B2 (en) * 2016-03-25 2020-03-10 General Electric Company Micro-channel cooling of integrated combustor nozzle of a segmented annular combustion system
US10605459B2 (en) * 2016-03-25 2020-03-31 General Electric Company Integrated combustor nozzle for a segmented annular combustion system
US11187415B2 (en) * 2017-12-11 2021-11-30 General Electric Company Fuel injection assemblies for axial fuel staging in gas turbine combustors
US11614233B2 (en) 2020-08-31 2023-03-28 General Electric Company Impingement panel support structure and method of manufacture
US11371702B2 (en) 2020-08-31 2022-06-28 General Electric Company Impingement panel for a turbomachine
US11460191B2 (en) 2020-08-31 2022-10-04 General Electric Company Cooling insert for a turbomachine
US11255545B1 (en) 2020-10-26 2022-02-22 General Electric Company Integrated combustion nozzle having a unified head end
US11846426B2 (en) * 2021-06-24 2023-12-19 General Electric Company Gas turbine combustor having secondary fuel nozzles with plural passages for injecting a diluent and a fuel
US20230228425A1 (en) * 2022-01-18 2023-07-20 Qingdao Zhennuo Laser Technology Co., Ltd. Multi-Nozzle Fuel Injection Method for Gas Turbine
US11898756B2 (en) * 2022-01-18 2024-02-13 Qingdao Zhennuo Laser Technology Co., Ltd. Multi-nozzle fuel injection method for gas turbine
US11767766B1 (en) 2022-07-29 2023-09-26 General Electric Company Turbomachine airfoil having impingement cooling passages
DE102022208339A1 (en) 2022-08-10 2024-02-15 Rolls-Royce Deutschland Ltd & Co Kg Combustion chamber arrangement for operation with liquid and/or gaseous fuel, gas turbine arrangement and method

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