WO2013028169A1 - Can-annular combustor with premixed tangential fuel-air nozzles for use on gas turbine engines - Google Patents

Can-annular combustor with premixed tangential fuel-air nozzles for use on gas turbine engines Download PDF

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Publication number
WO2013028169A1
WO2013028169A1 PCT/US2011/048622 US2011048622W WO2013028169A1 WO 2013028169 A1 WO2013028169 A1 WO 2013028169A1 US 2011048622 W US2011048622 W US 2011048622W WO 2013028169 A1 WO2013028169 A1 WO 2013028169A1
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WO
WIPO (PCT)
Prior art keywords
fuel
air
nozzles
annular combustor
combustor
Prior art date
Application number
PCT/US2011/048622
Other languages
French (fr)
Other versions
WO2013028169A8 (en
Inventor
Majed Toqan
Brent Allan Gregory
Jonathan David Regele
Ryan Sadao Yamane
Original Assignee
Majed Toqan
Brent Allan Gregory
Jonathan David Regele
Ryan Sadao Yamane
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 Majed Toqan, Brent Allan Gregory, Jonathan David Regele, Ryan Sadao Yamane filed Critical Majed Toqan
Priority to KR1020147007519A priority Critical patent/KR101774094B1/en
Priority to CN201180073014.0A priority patent/CN104053883B/en
Priority to EP11871108.4A priority patent/EP2748443B1/en
Priority to PCT/US2011/048622 priority patent/WO2013028169A1/en
Priority to JP2014527127A priority patent/JP6086371B2/en
Priority to RU2014110631A priority patent/RU2619673C2/en
Priority to PL11871108T priority patent/PL2748443T3/en
Publication of WO2013028169A1 publication Critical patent/WO2013028169A1/en
Publication of WO2013028169A8 publication Critical patent/WO2013028169A8/en

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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
    • F23R3/28—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
    • F23R3/42—Continuous combustion chambers using liquid or gaseous fuel characterised by the arrangement or form of the flame tubes or combustion chambers
    • F23R3/425—Combustion chambers comprising a tangential or helicoidal arrangement of the flame tubes
    • 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
    • F23R3/58—Cyclone or vortex type 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
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • 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
    • F23R3/30—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply comprising fuel prevapourising devices
    • 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

Definitions

  • This invention relates to devices in gas turbine engines that aid in containing and producing the combustion of a fuel and air mixture.
  • Such devices include but are not limited to fuel-air nozzles, combustor liners and casings and flow transition pieces that are used in military and commercial aircraft, power generation, and other gas turbine related applications.
  • Gas turbine engines include machinery that extracts work from combustion gases flowing at very high temperatures, pressures and velocity. The extracted work can be used to drive a generator for power generation or for providing the required thrust for an aircraft.
  • a typical gas turbine engine consists of a multistage compressor where the atmospheric air is compressed to high pressures. The compressed air is then mixed at a specified fuel/air ratio in a combustor wherein its temperature is increased. The high temperature and pressure combustion gases are then expanded through a turbine to extract work so as to provide the required thrust or drive a generator depending on the application.
  • the turbine includes at least a single stage with each stage consisting of a row of blades and a row of vanes. The blades are circumferentially distributed on a rotating hub with the height of each blade covering the hot gas flow path. Each stage of non-rotating vanes is placed circumferentially, which also extends across the hot gas flow path.
  • the included invention involves the combustor of gas turbine engines and components that introduce the fuel and air into the said device.
  • the combustor portion of a gas turbine engine can be of several different types: can/tubular, annular, and a combination of the two forming a can-annular combustor. It is in this component that the compressed fuel-air mixture passes through fuel-air swirlers and a combustion reaction of the mixture takes place, creating a hot gas flow causing it to drop in density and accelerate downstream.
  • the can type combustor typically comprises of individual, circumferentially spaced cans that contain the flame of each nozzle separately. Flow from each can is then directed through a duct and combined in an annular transition piece before it enters the first stage vane.
  • a fuel air nozzle can take on different configurations such as single to multiple annular inlets with swirling vanes on each one.
  • a typical method for cooling the combustor is effusion cooling, implemented by surrounding the combustion liner with an additional, offset liner, which between the two, compressor discharge air passes through and enters the hot gas flow path through dilution holes and cooling passages. This technique removes heat from the component as well as forms a thin boundary layer film of cool air between the liner and the combusting gases, preventing heat transfer to the liner.
  • the dilution holes serve two purposes depending on its axial position on the liner: a dilution hole closer to the fuel-air nozzles will aid in the mixing of the gases to enhance combustion as well as provide unburned air for combustion, second, a hole that is placed closer to the turbine will cool the hot gas flow and can be designed to manipulate the combustor outlet temperature profile.
  • the invention consists of a typical can-annular combustor with premixed fuel-air nozzles and/or dilution holes that introduce the compressor discharge air and pressurized fuel into the combustor at various locations in the longitudinal and circumferential directions.
  • the original feature of the invention is that the fuel and air nozzles are placed in such a way as to create an environment with enhanced mixing of combustion reactants and products.
  • Staging the premixed fuel and air nozzles to have more fuel upstream from another set of nozzles enhances the mixing of the combustion reactants and creates a specific oxygen concentration in the combustion region that greatly reduces the production of NOx.
  • the introduction of compressor discharge air downstream of the combustion region allows for any CO produced during combustion to be burned/consumed before entering the first stage turbine.
  • the combustor will improve gas turbine emission levels, thus reducing the need for emission control devices as well as minimize the environmental impact of such devices.
  • the tangentially firing fuel and fuel-air nozzles directs its flames to the adjacent burner nozzles in each can, greatly enhancing the ignition process of the combustor.
  • FIG. 1 is a two-dimensional sketch showing the can-annular arrangement with the nozzles that attach to the outer can liner injecting fuel and air into a common plane;
  • FIG. 2 is a two-dimensional sketch showing the general idea of the tangential nozzles applied to the can in a can-annular combustor;
  • FIG. 3 is an isometric side view of the upstream portion of an example configuration of the said invention.
  • FIG. 4A is an isometric cutaway view of the invention.
  • FIG. 4B is a close up view of the image from FIG. 4A;
  • FIG. 5 is a section view showing section A- A as defined in FIG. 3 ;
  • FIG. 6 is a section view showing section B-B as defined in FIG. 3. BEST MODES FOR CARRYING OUT THE INVENTION
  • FIG. 1 shows an example of the general arrangement of a can-annular combustor with the can 1 spaced circumferentially on a common radius, all cans of which are enclosed in an annular space between a cylindrical outer liner 2 and a cylindrical inner liner 3.
  • the FIG. also shows the tangential nozzle arrangement of the cans.
  • FIG. 2 shows the can in more detail.
  • a can liner 4 forms the can volume, with fuel-air nozzles 5 injecting a premixed fuel and air mixture.
  • the nozzles form an angle 8 between the nozzle centerline 6 and a line tangent to the can liner 4 that intersections with the nozzle centerline 6. This angle defines the circumferential direction of the nozzles.
  • FIG. 2 also shows the general operation of the can in the example can-annular combustor configuration, where a pre-mixed fuel-air mixture 9 is injected into the cans 1 at an angle 8.
  • a flame 10 forms and travels through the can in a path 11 that follows the can liner.
  • These tangentially directed nozzles result in flames from each nozzle interacting with the downstream and adjacent nozzle. This key feature enhances ignition and reduces the need of piloting burner nozzles by allowing the flame from a nozzle to ignite the fuel at the adjacent and downstream nozzle.
  • FIG. 3 shows the beginning or upstream portion of an example can with the downstream portion excluded.
  • the said invention will have a plurality of nozzle rows that are spaced along the longitudinal direction of the can.
  • Each row of nozzles may have at least one nozzle and can be offset by a circumferential angle from adjacent nozzle rows.
  • the can may also have several rows of circumferentially spaced holes 12 or passages for cooling air to enter the can at any location.
  • FIGS. 4A and 4B show the most upstream face 13 of the can, which may have holes 14 similar to dilution holes that allow compressor discharge air to enter the can.
  • FIGS. 5 and 6 show how nozzles from each set of rows may be offset by a circumferential angle. The different rows of nozzles allows for the injection of the fuel- air mixture near the front wall, which may have a higher fuel/air ratio than the second set of nozzles in conjunction with the mixture that is injected downstream of the fuel nozzles 5, to create the desired mixing and fuel- air staging effect that will create an optimal combustion environment that reduces NOx and CO emissions from the combustor.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Electrical Discharge Machining, Electrochemical Machining, And Combined Machining (AREA)

Abstract

A combustion device used in gas turbine engines to produce propulsion or rotate a shaft for power generation includes a can-annular combustor with a system of fuel and air inlet passages and nozzles that results in an optimal combustion environment of premixed fuel and air. The fuel-air inlets are placed at various longitudinal locations and circumferentially distributed, and direct the flow tangentially or nearly tangent to the can liner. The combustion device provides effective mixing of fuel and air, creates an environment for combustion that reduces pollutant emissions, reduces the need for costly pollution control devices, enhances ignition and flame stability, reduces piloting issues, and improves vibration reduction.

Description

CAN- ANNULAR COMBUSTOR WITH PREMIXED TANGENTIAL FUEL- AIR NOZZLES FOR USE ON GAS TURBINE ENGINES
TECHNICAL FIELD
This invention relates to devices in gas turbine engines that aid in containing and producing the combustion of a fuel and air mixture. Such devices include but are not limited to fuel-air nozzles, combustor liners and casings and flow transition pieces that are used in military and commercial aircraft, power generation, and other gas turbine related applications.
BACKGROUND ART
Gas turbine engines include machinery that extracts work from combustion gases flowing at very high temperatures, pressures and velocity. The extracted work can be used to drive a generator for power generation or for providing the required thrust for an aircraft. A typical gas turbine engine consists of a multistage compressor where the atmospheric air is compressed to high pressures. The compressed air is then mixed at a specified fuel/air ratio in a combustor wherein its temperature is increased. The high temperature and pressure combustion gases are then expanded through a turbine to extract work so as to provide the required thrust or drive a generator depending on the application. The turbine includes at least a single stage with each stage consisting of a row of blades and a row of vanes. The blades are circumferentially distributed on a rotating hub with the height of each blade covering the hot gas flow path. Each stage of non-rotating vanes is placed circumferentially, which also extends across the hot gas flow path. The included invention involves the combustor of gas turbine engines and components that introduce the fuel and air into the said device.
The combustor portion of a gas turbine engine can be of several different types: can/tubular, annular, and a combination of the two forming a can-annular combustor. It is in this component that the compressed fuel-air mixture passes through fuel-air swirlers and a combustion reaction of the mixture takes place, creating a hot gas flow causing it to drop in density and accelerate downstream. The can type combustor typically comprises of individual, circumferentially spaced cans that contain the flame of each nozzle separately. Flow from each can is then directed through a duct and combined in an annular transition piece before it enters the first stage vane. In the annular combustor type, fuel-air nozzles are typically distributed circumferentially and introduce the mixture into a single annular chamber where combustion takes place. Flow simply exits the downstream end of the annulus into the first stage turbine, without the need for a transition piece to combine the flow. The key difference of the last type, a can-annular combustor, is that it has individual cans encompassed by an annular casing that contains the air being fed into each can. Each variation has its benefits and disadvantages, depending on the application.
In combustors for gas turbines, it is typical for the fuel-air nozzle to introduce a swirl to the mixture for several reasons. One is to enhance mixing and thus combustion, another reason is that adding swirl stabilizes the flame to prevent flame blow out and it allows for leaner fuel-air mixtures for reduced emissions. A fuel air nozzle can take on different configurations such as single to multiple annular inlets with swirling vanes on each one.
As with other gas turbine components, implementation of cooling methods to prevent melting of the combustor material is needed. A typical method for cooling the combustor is effusion cooling, implemented by surrounding the combustion liner with an additional, offset liner, which between the two, compressor discharge air passes through and enters the hot gas flow path through dilution holes and cooling passages. This technique removes heat from the component as well as forms a thin boundary layer film of cool air between the liner and the combusting gases, preventing heat transfer to the liner. The dilution holes serve two purposes depending on its axial position on the liner: a dilution hole closer to the fuel-air nozzles will aid in the mixing of the gases to enhance combustion as well as provide unburned air for combustion, second, a hole that is placed closer to the turbine will cool the hot gas flow and can be designed to manipulate the combustor outlet temperature profile.
One can see that several methods and technologies can be incorporated into the design of combustors for gas turbine engines to improve combustion and lower emissions. While gas turbines tend to produce less pollution than other power generation methods, there is still room for improvement in this area. With government regulation of emissions tightening in several countries, the technology will need to improve to meet these requirements. DISCLOSURE OF THE INVENTION
With regard to present invention, there is provided a novel and improved combustor design that is capable of operating in a typical fashion while minimizing the pollutant emissions that are a result of combustion of a fuel and air mixture and address other issues faced by such devices. The invention consists of a typical can-annular combustor with premixed fuel-air nozzles and/or dilution holes that introduce the compressor discharge air and pressurized fuel into the combustor at various locations in the longitudinal and circumferential directions. The original feature of the invention is that the fuel and air nozzles are placed in such a way as to create an environment with enhanced mixing of combustion reactants and products. Staging the premixed fuel and air nozzles to have more fuel upstream from another set of nozzles enhances the mixing of the combustion reactants and creates a specific oxygen concentration in the combustion region that greatly reduces the production of NOx. In addition, the introduction of compressor discharge air downstream of the combustion region allows for any CO produced during combustion to be burned/consumed before entering the first stage turbine. In effect, the combustor will improve gas turbine emission levels, thus reducing the need for emission control devices as well as minimize the environmental impact of such devices. In addition to this improvement, the tangentially firing fuel and fuel-air nozzles directs its flames to the adjacent burner nozzles in each can, greatly enhancing the ignition process of the combustor.
BRIEF DESCRIPTION OF THE DRAWINGS
Referring to the drawings:
FIG. 1 is a two-dimensional sketch showing the can-annular arrangement with the nozzles that attach to the outer can liner injecting fuel and air into a common plane;
FIG. 2 is a two-dimensional sketch showing the general idea of the tangential nozzles applied to the can in a can-annular combustor;
FIG. 3 is an isometric side view of the upstream portion of an example configuration of the said invention;
FIG. 4A is an isometric cutaway view of the invention;
FIG. 4B is a close up view of the image from FIG. 4A;
FIG. 5 is a section view showing section A- A as defined in FIG. 3 ; and
FIG. 6 is a section view showing section B-B as defined in FIG. 3. BEST MODES FOR CARRYING OUT THE INVENTION
FIG. 1 shows an example of the general arrangement of a can-annular combustor with the can 1 spaced circumferentially on a common radius, all cans of which are enclosed in an annular space between a cylindrical outer liner 2 and a cylindrical inner liner 3. The FIG. also shows the tangential nozzle arrangement of the cans. FIG. 2 shows the can in more detail. A can liner 4 forms the can volume, with fuel-air nozzles 5 injecting a premixed fuel and air mixture. The nozzles form an angle 8 between the nozzle centerline 6 and a line tangent to the can liner 4 that intersections with the nozzle centerline 6. This angle defines the circumferential direction of the nozzles.
FIG. 2 also shows the general operation of the can in the example can-annular combustor configuration, where a pre-mixed fuel-air mixture 9 is injected into the cans 1 at an angle 8. A flame 10 forms and travels through the can in a path 11 that follows the can liner. These tangentially directed nozzles result in flames from each nozzle interacting with the downstream and adjacent nozzle. This key feature enhances ignition and reduces the need of piloting burner nozzles by allowing the flame from a nozzle to ignite the fuel at the adjacent and downstream nozzle.
FIG. 3 shows the beginning or upstream portion of an example can with the downstream portion excluded. The said invention will have a plurality of nozzle rows that are spaced along the longitudinal direction of the can. Each row of nozzles may have at least one nozzle and can be offset by a circumferential angle from adjacent nozzle rows. The can may also have several rows of circumferentially spaced holes 12 or passages for cooling air to enter the can at any location.
FIGS. 4A and 4B show the most upstream face 13 of the can, which may have holes 14 similar to dilution holes that allow compressor discharge air to enter the can. FIGS. 5 and 6 show how nozzles from each set of rows may be offset by a circumferential angle. The different rows of nozzles allows for the injection of the fuel- air mixture near the front wall, which may have a higher fuel/air ratio than the second set of nozzles in conjunction with the mixture that is injected downstream of the fuel nozzles 5, to create the desired mixing and fuel- air staging effect that will create an optimal combustion environment that reduces NOx and CO emissions from the combustor.
The present invention is described above with reference to a preferred embodiment. However, those skilled in the art will recognize that changes and modifications may be made in the described embodiment without departing from the nature and scope of the present invention. Various changes and modifications to the embodiment herein chosen for purposes of illustration will readily occur to those skilled in the art. To the extent that such modifications and variations do not depart from the spirit of the invention, they are intended to be included within the scope thereof.
Having fully described the invention in such clear and concise terms as to enable those skilled in the art to understand and practice the same, the invention claimed is:

Claims

1. A can-annular combustor for a gas turbine used in ground based power generation, land or sea based vehicles or aircraft engine applications, comprising: a plurality of circumferentially spaced can liners which are cylindrical in shape, each can having a plurality of tangentially pointing and circumferentially spaced fuel-air nozzles that share a common plane that is normal to the direction of the can centerline, with all liners made of high temperature alloys or a ceramic material.
2. The can-annular combustor as claimed in claim 1, wherein nozzles, circumferentially spaced in a common plane normal to the longitudinal direction that is near the front wall of the can injects a fuel-air mixture that has a greater fuel/air ratio than any downstream set of nozzles that may exist, and that mainly have a
circumferential direction and may have a radial and/or longitudinal direction.
3. The can-annular combustor as claimed in claim 1, wherein there may be nozzles, circumferentially spaced on at least one common plane normal to the longitudinal direction that is downstream from nozzles mentioned in claim 2, inject a fuel- air mixture that has a lower fuel/air ratio than that of the nozzles described in claim 2, and that mainly have a circumferential direction and may have a radial and/or longitudinal direction.
4. The can-annular combustor as claimed in claim 1 , wherein the nozzles may have constant or varying values of angle from plane to plane, as indicated by item 8 in FIG. 2, ranging from 0 to 90 degrees.
5. The can-annular combustor as claimed in claim 1, wherein the nozzles in the different planes may have the same fuel/air ratio or varying fuel/air ratio.
6. The can annular combustors as claimed in claim 1 , wherein the fuel air nozzles in the same plane may have the same fuel/air or varying values of fuel/air ratios.
7. The can- annular combustor as claimed in claim 1, wherein the tangentially directed nozzles greatly enhance the ignition process of the combustor because adjacent nozzle flames are pointed at the adjacent nozzle in its plane, which will reduce the need for piloting multiple burners.
8. The can-annular combustor as claimed in claim 5, wherein the enhanced ignition process produces inherently stable burners that will reduce flame induced vibrations and acoustics that are generated from flame instability at partial and full load operation.
9. The can-annular combustor as claimed in claim 1 , wherein the tangential fuel- air nozzle arrangement enhances mixing of reactants for efficient combustion at very low load levels.
10. The can-annular combustor as claimed in claim 1, wherein low reactivity fuels such as low BTU gases can be easily utilized and combusted in said combustor due to the increased flame stability.
11. The can-annular combustor as claimed in claim 1 , wherein a vortex is created about the can centerline (key result of tangential fuel-air nozzles) that promotes stable flames at the burner exit.
12. The can-annular combustor as claimed in claim 1, wherein the required residence time to combust the fuel-air mixture is reduced; as a result, combustion space is reduced, which decreases engine size (important in all applications stated) and thus weight to thrust ratio (important in aero gas turbine applications).
13. The can-annular combustor as claimed in claim 1, wherein a more uniform temperature distribution is achieved at the said combustor' s outlet which allows for it to operate at higher combustion (firing) temperatures without deteriorating the life of the combustor and turbine parts.
14. The can-annular combustor as claimed in claim 1, wherein the ability to operate at higher combustion temperature as stated in claim 13 results in increased engine efficiency and power output and thus reduces carbon dioxide emission levels.
15. The can- annular combustor as claimed in claim 1, wherein the can front wall liner may have at least one hole or nozzle that allows for compressor discharge air to penetrate said liner at velocity magnitude less than nozzles mentioned in claims 2,3.
16. The can-annular combustor as claimed in claim 1, wherein the radius and length of the cans may vary in the longitudinal direction depending upon the size and shape of the gas turbine engine.
17. The can-annular combustor as claimed in claim 1, wherein any cooling method available to cool gas turbine components may be used, for example: impingement cooling, effusion cooling, steam cooling, etc.
18. The can-annular combustor as claimed in claim 1, wherein the nozzles that share a common plane may be offset from another set of nozzles in a different plane by a circumferential angle about the can centerline.
19. The can-annular combustor as claimed in claim 1, wherein the air passages shown as items 12 and 14 may be a straight hole or have a bell mouth inlet, made using spark electrical discharge machining (EDM).
PCT/US2011/048622 2011-08-22 2011-08-22 Can-annular combustor with premixed tangential fuel-air nozzles for use on gas turbine engines WO2013028169A1 (en)

Priority Applications (7)

Application Number Priority Date Filing Date Title
KR1020147007519A KR101774094B1 (en) 2011-08-22 2011-08-22 Can-annular combustor with premixed tangential fuel-air nozzles for use on gas turbine engines
CN201180073014.0A CN104053883B (en) 2011-08-22 2011-08-22 Method for mixing combustion reactants combusting in gas turbine engine
EP11871108.4A EP2748443B1 (en) 2011-08-22 2011-08-22 Method of mixing combustion reactants for combustion in a gas turbine engine
PCT/US2011/048622 WO2013028169A1 (en) 2011-08-22 2011-08-22 Can-annular combustor with premixed tangential fuel-air nozzles for use on gas turbine engines
JP2014527127A JP6086371B2 (en) 2011-08-22 2011-08-22 Combustion reactant mixing method in annular cylindrical combustor for gas turbine engine
RU2014110631A RU2619673C2 (en) 2011-08-22 2011-08-22 Mixing of combustible substances procedure for gas turbine engine combustion chamber
PL11871108T PL2748443T3 (en) 2011-08-22 2011-08-22 Method of mixing combustion reactants for combustion in a gas turbine engine

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
PCT/US2011/048622 WO2013028169A1 (en) 2011-08-22 2011-08-22 Can-annular combustor with premixed tangential fuel-air nozzles for use on gas turbine engines

Publications (2)

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WO2013028169A1 true WO2013028169A1 (en) 2013-02-28
WO2013028169A8 WO2013028169A8 (en) 2014-04-17

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EP (1) EP2748443B1 (en)
JP (1) JP6086371B2 (en)
KR (1) KR101774094B1 (en)
CN (1) CN104053883B (en)
PL (1) PL2748443T3 (en)
RU (1) RU2619673C2 (en)
WO (1) WO2013028169A1 (en)

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CN107631323B (en) * 2017-09-05 2019-12-06 中国联合重型燃气轮机技术有限公司 Nozzle for gas turbine
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