EP3221643B1 - Combustion liner and method of reducing a recirculation zone of a combustion liner - Google Patents

Combustion liner and method of reducing a recirculation zone of a combustion liner Download PDF

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Publication number
EP3221643B1
EP3221643B1 EP15870609.3A EP15870609A EP3221643B1 EP 3221643 B1 EP3221643 B1 EP 3221643B1 EP 15870609 A EP15870609 A EP 15870609A EP 3221643 B1 EP3221643 B1 EP 3221643B1
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EP
European Patent Office
Prior art keywords
combustion liner
inlet end
coating
chamfer
combustion
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Active
Application number
EP15870609.3A
Other languages
German (de)
English (en)
French (fr)
Other versions
EP3221643A4 (en
EP3221643A2 (en
Inventor
Peter John Stuttaford
Hany Rizkalla
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Ansaldo Energia IP UK Ltd
Original Assignee
Ansaldo Energia IP UK Ltd
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
Priority claimed from US14/549,922 external-priority patent/US10060630B2/en
Application filed by Ansaldo Energia IP UK Ltd filed Critical Ansaldo Energia IP UK Ltd
Publication of EP3221643A2 publication Critical patent/EP3221643A2/en
Publication of EP3221643A4 publication Critical patent/EP3221643A4/en
Application granted granted Critical
Publication of EP3221643B1 publication Critical patent/EP3221643B1/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
    • F23R3/002—Wall structures
    • 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
    • F23R3/286—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply having fuel-air premixing 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
    • 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
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2240/00—Components
    • F05D2240/35—Combustors or associated equipment

Definitions

  • the present invention relates to a combustion liner and a method for reducing the recirculation zone in a combustion liner.
  • Diffusion type nozzles where fuel is mixed with air external to the fuel nozzle by diffusion, proximate the flame zone. Diffusion type nozzles historically produce relatively high emissions due to the fact that the fuel and air burn essentially upon interaction, without mixing, and stoichiometrically at high temperature to maintain adequate combustor stability and low combustion dynamics.
  • An alternate means of premixing fuel and air and obtaining lower emissions can occur by utilizing multiple combustion stages.
  • the fuel and air which mix and burn to form the hot combustion gases, must also be staged.
  • available power as well as emissions can be controlled.
  • Fuel can be staged through a series of valves within the fuel system or dedicated fuel circuits to specific fuel injectors.
  • Air can be more difficult to stage given the large quantity of air supplied by the engine compressor.
  • air flow to a combustor is typically controlled by the size of the openings in the combustion liner itself, and is therefore not readily adjustable.
  • FIG. 1 An example of the prior art combustion system 100 is shown in cross section in FIG. 1 .
  • the combustion system 100 includes a flow sleeve 102 containing a combustion liner 104.
  • a fuel injector 106 is secured to a casing 108 with the casing 108 encapsulating a radial mixer 110.
  • Secured to the forward portion of the casing 108 is a cover 112 and pilot nozzle assembly 114.
  • the present invention discloses a combustion liner and a method for reducing the recirculation zone in a combustion liner according to the attached set of claims.
  • the present invention discloses a system and method for controlling velocity of a fuel-air mixture being injected into a combustion system. That is, a predetermined effective flow area is maintained through two co-axial structures forming an annulus of a known effective flow area through which a fuel-air mixture passes.
  • FIG. 2 An embodiment of a gas turbine combustion system 200 in which the present invention operates is depicted in FIG. 2 .
  • the combustion system 200 is an example of a multi-stage combustion system and extends about a longitudinal axis A-A and includes a generally cylindrical flow sleeve 202 for directing a predetermined amount of compressor air along an outer surface of a generally cylindrical and co-axial combustion liner 204.
  • the combustion liner 204 has an inlet end 206 and opposing outlet end 208.
  • the combustion system 200 also comprises a set of main fuel injectors 210 that are positioned radially outward of the combustion liner 204 and proximate an upstream end of the flow sleeve 202.
  • the set of main fuel injectors 210 direct a controlled amount of fuel into the passing air stream to provide a fuel-air-mixture for the combustion system 200.
  • the main fuel injectors 210 are located radially outward of the combustion liner 204 and spread in an annular array about the combustion liner 204.
  • the main fuel injectors 210 are divided into two stages with a first stage extending approximately 120 degrees about the combustion liner 204 and a second stage extending the remaining annular portion, or approximately 240 degrees, about the combustion liner 204.
  • the first stage of the main fuel injectors 210 are used to generate a Main 1 flame while the second stage of the main fuel injectors 210 generate a Main 2 flame.
  • the combustion system 200 also comprises a combustor dome assembly 212, which, as shown in FIGS. 2 and 3 , encompasses the inlet end 206 of the combustion liner 204. More specifically, the dome assembly 212 has an outer annular wall 214 that extends from proximate the set of main fuel injectors 210 to a generally hemispherical-shaped cap 216, which is positioned a distance forward of the inlet end 206 of the combustion liner 204. The dome assembly 212 turns through the hemispherical-shaped cap 216 and extends a distance into the combustion liner 204 through a dome assembly inner wall 218.
  • a first passageway 220 is formed between the outer annular wall 214 and the combustion liner 204.
  • a first passageway 220 tapers in size, from a first radial height H1 proximate the set of main fuel injectors 210 to a smaller height H2 at a second passageway 222.
  • the first passageway 220 tapers at an angle to accelerate the flow to a target threshold velocity at a location H2 to provide adequate flashback margin. That is, when velocity of a fuel-air mixture is high enough, should a flashback occur in the combustion system, the velocity of the fuel-air mixture through the second passageway will prevent a flame from being maintained in this region.
  • the second passageway 222 is formed between a cylindrical portion of the outer annular wall 214 and the combustion liner 204, proximate the inlet end 206 of the combustion liner and is in fluid communication with the first passageway 220.
  • the second passageway 222 is formed between two cylindrical portions and has a second radial height H2 measured between the outer surface of the combustion liner 204 and the inner surface of the outer annular wall 214.
  • the combustor dome assembly 212 also comprises a third passageway 224 that is also cylindrical and positioned between the combustion liner 204 and the outer annular wall 214.
  • the combustor dome assembly 212 also comprises a third passageway 224 that is also cylindrical and positioned between the combustion liner 204 and inner wall 218.
  • the third passageway has a third radial height H3, and like the second passageway, is formed by two cylindrical walls - combustion liner 204 and dome assembly inner wall 218.
  • the first passageway 220 tapers into the second passageway 222, which is generally cylindrical in nature.
  • the second radial height H2 serves as the limiting region through which the fuel-air mixture must pass.
  • the radial height H2 is regulated and kept consistent from part-to-part by virtue of its geometry, as it is controlled by two cylindrical (i.e. not tapered) surfaces, as shown in FIG. 3 . That is, by utilizing a cylindrical surface as a limiting flow area, better dimensional control is provided because more accurate machining techniques and control of machining tolerances of a cylindrical surface is achievable, compared to that of tapered surfaces. For example, it is well within standard machining capability to hold tolerances of cylindrical surfaces to within +/- 0,00254 cm (0.001 inches).
  • Utilizing the cylindrical geometry of the second passageway 222 and third passageway 224 provides a more effective way to control and regulate the effective flow area and controlling the effective flow area allows for the fuel-air mixture to be maintained at predetermined and known velocities. By being able to regulate the velocity of the mixture, the velocity can be maintained at a rate high enough to ensure flashback of the flame does not occur in the dome assembly 212.
  • One such way to express these critical passageway geometries shown in FIGS. 2-4B is through a turning radius ratio of the second passageway height H2 relative to the third passageway height H3. That is, the minimal height relative to the height of the combustion inlet region.
  • the ratio of H2/H3 is approximately 0.32.
  • This aspect ratio controls the size of the recirculation and stabilization trapped vortex that resides adjacent to the liner, which effects overall combustor stability.
  • utilizing this geometry permits velocity of the fuel-air mixture in the second passageway to remain within a range of approximately 40-80 meters per second.
  • the ratio can vary depending on the desired passageway heights, fuel-air mixture mass flow rate and combustor velocities.
  • the ratio of H2/H3 can range from approximately 0.1 to approximately 0.5. More specifically, for an embodiment of the present invention, the first radial height H1 can range from approximately 15 millimeters to approximately 50 millimeters, while the second radial height H2 can range from approximately 10 millimeters to approximately 45 millimeters, and the third radial height H3 can range from approximately 30 millimeters to approximately 100 millimeters.
  • the combustion system also comprises a fourth passageway 226 having a fourth height H4, where the fourth passageway 226 is located between the inlet end 206 of the combustion liner and the hemispherical-shaped cap 216.
  • the fourth passageway 226 is positioned within the hemispherical-shaped cap 216 with the fourth height measured along the distance from the inlet end 206 of the liner to the intersecting location at the hemispherical-shaped cap 216.
  • the fourth height H4 is greater than the second radial height H2, but the fourth height H4 is less than the third radial height H3.
  • This relative height configuration of the second, third and fourth passageways permits the fuel-air mixture to be controlled (at H2), turn through the hemispherical-shaped cap 216 (at H4) and enter the combustion liner 204 (at H3) all in a manner so as to ensure the fuel-air mixture velocity is fast enough that the fuel-air mixture remains attached to the surface of the dome assembly 212, as an unattached, or separated, fuel-air mixture could present a possible condition for supporting a flame in the event of a flashback.
  • the height of the first passageway 220 tapers as a result, at least in part, of the shape of outer annular wall 214. More specifically, the first passageway 220 has its largest height at a region adjacent the set of main fuel injectors 210 and its minimum height at the region adjacent the second passageway. Alternate embodiments of the dome cap assembly 212 having the passageway geometry described above are shown in better detail in FIGS. 4A and 4B .
  • a method 500 of controlling a velocity of a fuel-air mixture for a gas turbine combustor comprises a step 502 of directing a fuel-air mixture through a first passageway that is located radially outward of a combustion liner. Then, in a step 504, the fuel-air mixture is directed from the first passageway and into a second passageway that is also located radially outward of the combustion liner. In a step 506, the fuel-air mixture is directed from the second passageway and into the fourth passageway formed by the hemispherical dome cap 216. As a result, the fuel-air mixture reverses its flow direction to now be directed into the combustion liner. Then, in a step 508, the fuel-air mixture is directed through a third passageway located within the combustion liner such that the fuel-air mixture passes downstream into the combustion liner.
  • a gas turbine engine typically incorporates a plurality of combustors.
  • the gas turbine engine may include low emission combustors such as those disclosed herein and may be arranged in a can-annular configuration about the gas turbine engine.
  • One type of gas turbine engine e.g., heavy duty gas turbine engines
  • the combustion system 200 disclosed in FIGS. 2 and 3 is a multi-stage premixing combustion system comprising four stages of fuel injection based on the loading of the engine.
  • the specific fuel circuitry and associated control mechanisms could be modified to include fewer or additional fuel circuits.
  • FIG. 6 a detailed view of the inlet end of a combustion liner of the prior art is shown. More specifically, a combustion liner 600 has a generally annular body 602 with a thickness 604 and a thermal barrier coating 606 applied along an inner surface 608 of the generally annular body 602. The combustion liner 600 has an inlet end 610. In this prior art embodiment, the thermal barrier coating 606 extends to the inlet end 610, and together forms a blunt face 612.
  • the inlet end 610 has a combined thickness (metal + thermal barrier coating) upwards of 0,2286 cm (0.090 inches) or greater, depending on the sheet metal thickness used for the combustion liner 600.
  • a combustion liner 600 is used in conjunction with a combustion system of FIGS. 2-5 , the combustion liner 600 and its inlet end 610 form a bluff body that can yield undesirable results when the flow of fuel and air pass along and around the inlet end 610. More specifically, as the flow of fuel and air pass around the inlet end 610, the fuel and air mixture tends to separate as it enters the combustion liner 600 due to the bluff body geometry.
  • flow separation such as this can help to anchor a flame at or near the inlet end 610.
  • This undesirable result causes the inlet end 610 of the combustion liner 600 to be eroded by the flame formed in this area of recirculation resulting in premature repair or replacement to the combustion liner.
  • a combustion liner 700 having a generally annular body 702 having a thickness T that varies towards a forward region 704.
  • the combustion liner 700 also has an inlet end 706 and an opposing outlet end (not shown).
  • the generally annular body 702 also has an inner surface 708 and an opposing outer surface having a contoured profile proximate the inlet end 706 comprising a first outer surface 710 and a second outer surface 712 where the first outer surface 710 is located radially outward of the second outer surface 712.
  • the forward region 704 of the combustion liner 700 also has a first chamfer 714 extending from the first outer surface 710 towards the inlet end 706, thereby reducing the thickness of the combustion liner 700 in the forward region 704.
  • the first chamfer 714 is oriented at approximately a 5-75 degree angle and reduces the thickness of the combustion liner 700 from approximately 0,254 - 0,635 cm (0.1-0.25 inches) to approximately 0,0127- 0,254 cm (0.005 - 0.1 inches) at the inlet end 706.
  • the chamfer angle, resulting thickness, and rate of change for the thickness of the combustion liner are merely representative and not meant to be limiting the scope of the present invention.
  • the thickness of the combustion liner, chamfer angle, and rate of thickness change towards the inlet end 706 can vary. However, by tapering the thickness change via first chamfer 714 at a first rate, more of the flow of fuel and air passing along the outer surface of the generally annular body 702 remains attached to the annular body 702 as opposed to prior art designs.
  • the combustion liner 700 also comprises a coating 716 applied to the inner surface 708 of the generally annular body 702.
  • a coating utilized for the combustion liner 700 is a thermal barrier coating.
  • the thermal barrier coating 716 applied to the inner surface 708 comprises a bond coating 718 and a ceramic top coating 720.
  • the bond coating 718 can be applied approximately 0,00254 - 0,0254 cm (0.001 - 0.010 inches) thick
  • the ceramic top coating 720 can be applied approximately 0,0254-0,508 cm (0.010 - 0.200 inches) thick over the bond coating 718.
  • the thermal barrier coating can be a standard commercial coating discussed above or can also be a more advanced thermal barrier coating such as a dense vertically cracked coating.
  • a portion of the coating proximate the inlet end 706 is tapered via a second chamfer 722 oriented at an angle of 5-75 degrees, which tapers the coating thickness towards the inlet end 706 at a second rate.
  • the second chamfer 722 can be formed via a machining process, such as grinding to a previously-applied coating, or it can be formed as a result of tapering the layers of bond coating and thermal barrier coating applied.
  • the first chamfer 714 and the second chamfer 722 form a reduced bluff body region 724 at the inlet end 706.
  • the reduced bluff body region 724 has a thickness of approximately 0,0508 cm (0.020 inches).
  • other reduced bluff body regions 724 can be utilized depending on the desired configuration of the combustion liner 700.
  • a bluff body region creates a recirculation zone.
  • the chamfer angles 714 and 722 of the present invention reduce the size of such a region so as to reduce the tendency for the flow of fuel and air to separate as it passes towards the inlet end 706.
  • the flow of fuel and air passing along the outer region of the generally annular body 702 remains along the tapered surfaces 714 and 722, thereby reducing the adverse effect of the bluff body of the prior art.
  • the chamfer at the liner inlet end 706 may instead comprise a rounded bluff body region or a rounded portion of the liner inlet end as shown in FIGS. 8-10 .
  • a combustion liner 800 has an inlet end 806 and instead of the chamfer angles 714 and 722 shown in FIG. 7 , the combustion liner 800 has one or more radii at the inlet end 806. That is, the combustion liner 800 comprises a generally annular body 802 with an inlet end 806 and an outlet end (not shown).
  • the annular body 802 has an inner surface 808 and an outer surface 810. In this embodiment, the inner surface 808 has a thermal barrier coating 820 applied thereto.
  • this embodiment includes one or more radii formed into the combustion liner 800 at the liner inlet. More specifically, in FIG. 8 , which is not part of the present invention, the one or more radii comprise a radius R to the generally annular body 802 about the outer surface 810 proximate inlet end 806. Radius R can vary depending on a variety of factors. However, it is preferred that radius R extends a distance so as to extend generally equivalent to the length of the tapered surface 714 of the embodiment in FIG. 7 . As such, the radius R covers the same general region of the tapered surface 714. However, while a radius provides a similar benefit to that of the tapered surface 714, it is not as advantageous as the tapered surface 714. The radius R increases the risk of separation of the air flow as a result of the curved surface. Also, such a radius negatively affects any flame holding in the area.
  • the one or more radius R to the combustion liner 800 is formed along the thermal barrier coating 820 applied to the inner surface 808 at the inlet end 806.
  • the radius R of the thermal barrier coating 820 can vary depending on the coating thickness. As with the embodiment of FIG. 8 , the radius R to the thermal barrier coating also negatively affects flame holding in the inlet end 806.
  • the one or more radii R can comprise a first radius R1 and a second radius R2. More specifically, the generally annular body 802 has a first radius R1 that is generally greater than the radius R2 of the thermal barrier coating 820. As such, the combination of R1 and R2 at the inlet end 806 forms a shape comparable to a bullnose at the inlet to the combustion liner.
  • FIGS. 8-10 provide a blunt front edge of the combustion liner that is necessary for the liner structural integrity. However, reducing the front edge thickness prevents premature thermal wear of the combustion liner inlet end 806 by reducing the tendency for flame holding.
  • the radii R, R1 and/or R2 are formed preferably by a grinding process to the liner and/or thermal barrier coating.
  • a method 1100 of reducing a recirculation zone in a gas turbine combustor is disclosed. More specifically, in a step 1102, a combustion liner is provided having a chamfer along an outer surface of the combustion liner, a coating applied to an inner surface of the combustion liner, and a chamfer to the coating on the inner surface. Then, in a step 1104, a fuel and air mixture is directed along the outer surface of the combustion liner. The fuel and air mixture is then turned about an inlet end of the combustion liner in a step 1106, such that the mixture remains at least in close proximity to the chamfered portions of the combustion liner. Then, in a step 1108, the fuel and air mixture is directed into the combustion liner where it is ignited to supply power to the gas turbine engine.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Coating By Spraying Or Casting (AREA)
EP15870609.3A 2014-11-21 2015-11-20 Combustion liner and method of reducing a recirculation zone of a combustion liner Active EP3221643B1 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
US14/549,922 US10060630B2 (en) 2012-10-01 2014-11-21 Flamesheet combustor contoured liner
PCT/US2015/061926 WO2016099805A2 (en) 2014-11-21 2015-11-20 Flamesheet combustor contoured liner

Publications (3)

Publication Number Publication Date
EP3221643A2 EP3221643A2 (en) 2017-09-27
EP3221643A4 EP3221643A4 (en) 2018-09-05
EP3221643B1 true EP3221643B1 (en) 2020-02-26

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EP15870609.3A Active EP3221643B1 (en) 2014-11-21 2015-11-20 Combustion liner and method of reducing a recirculation zone of a combustion liner

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EP (1) EP3221643B1 (ko)
JP (1) JP6824165B2 (ko)
KR (1) KR102405991B1 (ko)
CN (1) CN107429920B (ko)
WO (1) WO2016099805A2 (ko)

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US11859819B2 (en) 2021-10-15 2024-01-02 General Electric Company Ceramic composite combustor dome and liners

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US10684014B2 (en) 2016-08-04 2020-06-16 Raytheon Technologies Corporation Combustor panel for gas turbine engine
US10823410B2 (en) 2016-10-26 2020-11-03 Raytheon Technologies Corporation Cast combustor liner panel radius for gas turbine engine combustor

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Publication number Publication date
JP6824165B2 (ja) 2021-02-03
CN107429920B (zh) 2019-11-05
JP2017536524A (ja) 2017-12-07
KR102405991B1 (ko) 2022-06-10
WO2016099805A3 (en) 2016-10-27
EP3221643A4 (en) 2018-09-05
EP3221643A2 (en) 2017-09-27
WO2016099805A2 (en) 2016-06-23
KR20170106302A (ko) 2017-09-20
CN107429920A (zh) 2017-12-01

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