CN107429920B - Flame front burner determines the bushing of shape - Google Patents
Flame front burner determines the bushing of shape Download PDFInfo
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- CN107429920B CN107429920B CN201580074243.2A CN201580074243A CN107429920B CN 107429920 B CN107429920 B CN 107429920B CN 201580074243 A CN201580074243 A CN 201580074243A CN 107429920 B CN107429920 B CN 107429920B
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- coating
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- combustion
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Classifications
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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
- F23R3/002—Wall structures
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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
- F23R3/286—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply having fuel-air premixing devices
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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
- F23R3/42—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
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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
- F23R3/42—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
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- 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
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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)
Abstract
The novel device and mode for the recirculation regions at arrival end that the invention discloses a kind of for reducing burner.Tapered by the tapered and thermal barrier coating via wall thickness changes the geometry of arrival end to reduce the blunt body effect at combustion liner arrival end, reduces recirculation regions.
Description
Technical field
The present invention relates generally to a kind of equipment and sides for being directed to fuel-air mixture in combustion system
Method.More specifically, hemispherical dome is located in the entrance of combustion liner to guide Fuel-air in a more effective manner
Mixture enters the speed of combustion liner preferably to control fuel-air mixture, while making at combustion liner entrance area
Unfavorable aerodynamic effect it is as small as possible.
Background technique
In order to reduce the discharge amount of pollution of the turbine from gas energy supply, government bodies have formulated requirement and have reduced nitrogen oxides
(NOx) and many regulations of the amount of carbon monoxide (CO).Lower burning and exhausting is generally attributable to more effective combustion process,
Especially in regard to fuel injector position, air stream rate and mixing effectiveness.
Earlier combustion system use diffusion types nozzle, wherein fuel by the diffusion near flame region come with combustion
Expect the air mixing outside nozzle.The nozzle of diffusion types generates relatively high discharge in the past, because of in fact, fuel and air
It (does not mix) in interaction substantially and is stoichiometrically burned under high-temperature, to maintain enough burner stables
It is qualitative and low combustion powered.
So that fuel and air is premixed and is obtained can be occurred compared with the alternative means of low emission by using multiple combustion stages.
In order to provide multistage combustion to burner, mixing and burning must be also classified to form the fuel and air of hot combustion gas.It is logical
The amount for the fuel and air that control is transmitted in combustion system is crossed, can control available power and discharge.Fuel is via fuel system
A series of valve in system or the special fuel route classification to special fuel injector.However, it is assumed that large quantity of air is by starting
The supply of machine compressor, then air may be more difficult to be classified.In fact, due to the general design of gas turbine combustion system, such as by Fig. 1
It is shown, until the air stream of burner is usually controlled by the size of the opening in combustion liner itself, and is therefore not easily adjusted.
The example of the combustion system 100 of the prior art is shown in Fig. 1 with section.Combustion system 100 includes comprising combustion liner 104
Flow sleeve 102.The dress of fuel injector 106 is affixed on shell 108, and wherein shell 108 encapsulates radial mixer 110.Lid 112 and guidance
The dress of nozzle assembly 114 is affixed on the front part of shell 108.
However, although before combustion having been shown fuel and air premix facilitates compared with low emission, the combustion of injection
Material-air pre-mixing object amount modified trend due to a variety of burner variables.Burner is ejected into accordingly, with respect to control
In fuel-air premixed material amount, there are still obstacles.
Summary of the invention
The invention discloses a kind of for improving before in the combustion liner that mixture is ejected into multi-stage combustion system
The device and method of the control of Fuel-air mixing.More specifically, in an embodiment of the present invention, gas turbine combustor
It is set as with generally columnar flow sleeve and the generally columnar combustion liner being included in.Gas turbine combustor
It further include the arrival end of one group of main fuel injector and encirclement combustion liner and the burner with generally hemispherical cross-section
Dome component.Dome component axially extends both towards this group of main fuel injector and in combustion liner, to form fuel-
Air mixture is conveyed through a series of accesses therebetween, and wherein access is correspondingly really sized to adjust Fuel-air premix
The stream of object.
In alternative embodiment of the invention, a kind of dome component for gas turbine combustor is disclosed.Dome group
Part includes the outer ring on the radially outer point that hemispherical cover, the dress of the annular that the axis of burner extends are affixed to hemispherical cover
The inner annular wall that shape wall and also dress are affixed in the radially inner portion of hemispherical cover.Resulting dome component is determined with size
At the essentially U-shaped section for the intake section for surrounding combustion liner.
In yet another embodiment of the present invention, a kind of Fuel-air for controlling and being used for gas turbine combustor is disclosed
The method of the speed of mixture.This method includes the radial outside for guiding fuel-air mixture to be placed through combustion liner
First access, and fuel-air mixture is then directed across second leading near the first access from the first access
Road.Then fuel-air mixture guides from alternate path and passes through the fourth passage formed by hemispherical dome cover, to draw
Play fuel-air mixture reverses direction.It is logical that fuel-air mixture is then passed through the third in combustion liner
Road.
In yet another embodiment of the present invention, substantially ring-like ontology is provided, with thickness, arrival end, phase
Pair outlet end, inner surface and opposite outer surface, wherein outer surface has the profile for determining shape near arrival end so that
Outer surface includes the first outer surface and the second outer surface, wherein the first outer surface is located at the radial outside of the second outer surface, and the
One chamfering extends to arrival end from the first outer surface.Thermal barrier coating is applied on inner surface, floating coat near arrival end
A part have the second chamfering, to keep coating layer thickness tapered towards arrival end.
In another embodiment of the present invention, the intake section of combustion liner is provided comprising from the first bushing thickness
Tapered substantially ring-like ontology is spent, which has the second insert thickness, and from the first insert thickness near arrival end
It is tapered with first rate.Coating is applied on the inner wall of substantially ring-like ontology, coating from first coating thickness be tapered into
Second coating thickness at mouth end, coating are tapered with the second rate.
In yet another embodiment of the present invention, a kind of method for reducing the recirculation regions in combustion liner is provided.
Combustion liner be set as with along the outer surface of combustion liner chamfering, be applied to combustion liner inner surface on coating and
The chamfering of coating on inner surface.The mixture of fuel and air is guided along the outer surface of combustion liner, and around burning lining
Close to the chamfered part of combustion liner and then the arrival end of set turns to, so that mixture keeps at least being directed to combustion liner
In.
In another alternative embodiment of the invention, combustion liner is set as including substantially ring-like ontology, has
Thickness, arrival end, opposite outlet end, inner surface and opposite outer surface, wherein outer surface has the fixed outer of the first radius of band
The profile of shape.Thermal barrier coating is applied on inner surface, and a part near arrival end of floating coat has chamfering, to make
The arrival end of coating layer thickness towards combustion liner is tapered.
In another alternative embodiment of the invention, combustion liner is set as including substantially ring-like ontology, has
Thickness, arrival end, opposite outlet end, inner surface and opposite outer surface, wherein outer surface has the entrance towards combustion liner
The chamfering profile at end.Thermal barrier coating is applied on inner surface, and a part near arrival end of floating coat has band first
Radius determines the profile of shape, to keep coating layer thickness tapered towards the arrival end of combustion liner.
In another alternative embodiment of the invention, combustion liner is set as including substantially ring-like ontology, has
Thickness, arrival end, opposite outlet end, inner surface and opposite outer surface, wherein outer surface has the fixed outer of the first radius of band
The profile of shape.Thermal barrier coating is applied on inner surface, and a part near arrival end of floating coat has the second radius, from
And keep coating layer thickness tapered towards arrival end.
Additional advantage and feature of the invention will be set forth in part in the description that follows, and when consulting following for ability
The technical staff in domain will become obvious, or can learn from implementation of the invention to.It will be retouched now referring in particular to attached drawing
State the present invention.
Detailed description of the invention
Below with reference to attached drawing the present invention is described in detail, in the accompanying drawings:
Fig. 1 is the section of the combustion system of the prior art.
Fig. 2 is the section of the gas turbine combustor of embodiment according to the present invention.
Fig. 3 is the detail section of a part of the gas turbine combustor of Fig. 2 of embodiment according to the present invention.
Fig. 4 A is the section view of the dome component of embodiment according to the present invention.
Fig. 4 B is the section view of the dome component of alternative embodiment according to the present invention.
Fig. 5 is to disclose the flow chart for the process for adjusting the fuel-air mixture for entering gas turbine combustor.
Fig. 6 is the section view of a part of combustion liner according to prior art.
Fig. 7 is the section view of a part of the combustion liner of embodiment according to the present invention.
Fig. 8 is the section view of a part of the combustion liner of alternative embodiment according to the present invention.
Fig. 9 is the section view of a part of the combustion liner of another alternative embodiment according to the present invention.
Figure 10 is the section view of a part of combustion liner according to another embodiment of the invention.
Figure 11 is to depict embodiment according to the present invention the mixture of fuel and air is directed in combustion liner
Process flow chart.
Specific embodiment
By reference, the application combine U.S. Patent No. 6935116,6986254,7137256,7237384,
7308793,7513115 and No. 7677025 themes.
The invention discloses a kind of systems for controlling the speed for the fuel-air mixture being ejected into combustion system
And method.That is, by formed fuel-air mixture be conveyed through known to effective flow region two of annulus coaxial knots
Structure maintenance makes a reservation for effective flow region.
The present invention will be discussed about Fig. 2-8 now.The present invention is depicted in Fig. 2 in the gas turbine combustion wherein operated
The embodiment of system 200.Combustion system 200 is the example of multi-stage combustion system, and is extended around longitudinal axis A-A, and including
Generally columnar flow sleeve 202 is for predetermined along the guidance of the generally outer surface of column and coaxial combustion liner 204
The compressor air of amount.Combustion liner 204 has arrival end 206 and opposite outlet end 208.Combustion system 200 further includes one
Group main fuel injector 210, is located in the radial outside of combustion liner 204 and near the upstream end of flow sleeve 202.It should
The fuel of controlled quatity is directed in the air stream of transmission by group main fuel injector 210, to provide the combustion for being used for combustion system 200
Material-air mixture.
For the embodiment of the present invention shown in Fig. 2, the diameter that main fuel injector 210 is located at combustion liner 204 is outside
Side, and spread around combustion liner 204 with annular array.Main fuel injector 210 is divided into two-stage, and wherein the first order is around burning
Bushing 204 extends 120 degree of approximation, and the second level extends remaining annular section (or 240 degree approximate) around combustion liner 204.
The first order of main fuel injector 210 is for generating main 1 flame, and the second level of main fuel injector 210 generates main 2 flames.
Combustion system 200 further includes combustor dome component 212, surrounds combustion liner as in figs. 2 and 3
204 arrival end 206.More specifically, dome component 212 has annular wall 214, from this group of main fuel injector 210
It extends about to generally hemispherical cover 216 (at a certain distance from its 206 front of arrival end for being located in combustion liner 204).
Dome component 212 is turned to by hemispherical cover 216, and extends certain distance to combustion liner 204 by dome component inner wall 218
In.
Due to the geometry of the combustor dome component 212 together with combustion liner 204, a series of accesses are formed in burning
Between the part and combustion liner 204 of device dome component 212.First access 220 is formed in annular wall 214 and combustion liner
Between 204.Referring to Fig. 3, the first access 220 by size from the first radial height H1 near this group of main fuel injector 210 gradually
The relatively low height H2 being reduced at alternate path 222.Angularly tapered will flow accelerates to target at the H2 of position to first access 220
Threshold velocity, to provide enough tempering (flashback) nargin.That is, when the speed of fuel-air mixture is sufficiently high,
If be tempered in combustion system, the speed that fuel-air mixture passes through alternate path will prevent flame from maintaining
In the region.
Alternate path 222 is formed between the stylolitic part of annular wall 214 and combustion liner 204, in combustion liner
Arrival end 206 near, and with the first access 220 be in fluid communication.Alternate path 222 is formed between two stylolitic parts, and
With the second radial height H2 measured between the outer surface of combustion liner 204 and the inner surface of annular wall 214.Burner
Dome component 212 further includes third path 224, is also columnar, and is located between combustion liner 204 and inner wall 218.The
Three-way has third radial height H3, and by two cylindrical walls (in combustion liner 204 and dome component as alternate path
Wall 218) it is formed.
As discussed before, the first access 220 tapers to alternate path 222 (it is substantially columnar in nature)
In.Second radial height H2 is used as the restricted area that fuel-air mixture must be conveyed through.As shown in Figure 3, radial high
Degree H2 by means of its geometry adjust with remain from part to part it is consistent because it is by two columns (that is, not tapered)
Granule surface contral.That is, by the way that cylinder surface with stream region is restricted, is provided preferable size Control, because compared to tapered
, it can be achieved that the more acurrate machining technique of cylinder surface and the control of machining tolerances for surface.For example, by cylinder surface
It is sufficiently in Standard Machine working ability that tolerance, which is maintained in +/- 0.001 inch,.
Provide more efficient way using the cylindrical geometry of alternate path 222 and third path 224 controlling and
Effective flowing region is adjusted, and control effective flowing region allows for fuel-air mixture and maintains predetermined and known speed
Degree.By that can adjust the speed of mixture, speed can be maintained high enough to ensuring that the tempering of flame will not be in dome component
The rate occurred in 212.
A kind of such mode for expressing these critical path geometries shown in Fig. 2-4B is by alternate path height
Spend turning radius ratio of the H2 relative to third path height H3.That is, the minimum constructive height of the height relative to combustion inlet region.Example
Such as, in the embodiment of the present invention shown in this article, the ratio between H2/H3 is approximation 0.32.Length-width ratio control is located near bushing
Recycling and stablize trapping vortex size, this influence overall burner stability.For example, for real shown in Fig. 2 and 3
Example is applied, the speed that the fuel-air mixture in alternate path is allowed for using the geometry is maintained at approximate 40-80 per second
In the range of rice.However, this is than may depend on desired passage in height, fuel-air mixture mass flowrate and burner speed
It spends and changes.For disclosed combustion system, the ratio between H2/H3 can range from approximation 0.1 to approximation 0.5.More particularly, for
The embodiment of the present invention, the first radial height H1 can range from approximate 15 millimeters to 50 millimeters approximate, and the second radial height H2
Can range from approximate 10 millimeters to 45 millimeters approximate, and third radial height H3 can range from approximate 30 millimeters to 100 milli of approximation
Rice.
As discussed before, combustion system further includes the fourth passage 226 with the 4th height H4, wherein four-way
Road 226 is located between the arrival end 206 and hemispherical cover 216 of combustion liner.As it can be seen in Fig. 3, fourth passage 226 is located in half
In enclosed globe shade 216, wherein the distance of intersection position of the 4th height along the arrival end 206 from bushing to from hemispherical cover 216 is surveyed
.Therefore, the 4th height H4 is greater than the second radial height H2, but the 4th height H4 is less than third radial height H3.Alternate path,
This of third path and fourth passage relative altitude construction allow for fuel-air mixture by control (at H2), and steering is worn
More than half enclosed globe shades 216 (at H4), and combustion liner 204 (at H3) is fully entered in one way, to ensure fuel-sky
Gas mixture speed is sufficiently fast, so that fuel-air mixture remains attached on the surface of dome component 212, because not attached
Or isolated fuel-air mixture the possible condition of flame support will be brought in tempering.
As it can be seen in Fig. 3, the height of the first access 220 at least partially due to the shape of annular wall 214 and it is tapered.More
Specifically, the first access 220 has its maximum height at the region near this group of main fuel injector 210, and second
There is its minimum constructive height at region near access.It is illustrated in greater detail in Fig. 4 A and 4B with above-described access geometry
The alternative embodiment of the dome shade assembly 212 of shape.
Fig. 5 is turned to, method 500 of the control for the speed of the fuel-air mixture of gas turbine combustor is disclosed.
Method 500 includes the steps that the first access 502 of the radial outside for guiding fuel-air mixture to be placed through combustion liner.
Then, in step 504, fuel-air mixture guides from the first access and into alternate path, and alternate path also is located at combustion
Burn the radial outside of bushing.In step 506, fuel-air mixture guides from alternate path and to by hemispherical dome cover
In 216 fourth passages formed.As a result, fuel-air mixture reverses its flow direction and is being directed to combustion liner now
In.Then, in step 508, fuel-air mixture is directed across the third path in combustion liner, so that fuel-
Air mixture is delivered downstream in combustion liner.
As understood by those skilled in the art, gas-turbine unit usually combines multiple burners.Generally,
For the purpose of discussion, gas-turbine unit may include low emission combustor, such as those disclosed herein, and may be disposed to
Around tank-circular structure of gas-turbine unit.A type of gas-turbine unit is (for example, heavily loaded gas turbine hair
Motivation) it can be typically provided with but be not limited to six to ten eight individual burners, each of which is matched with component proposed above
It closes.Therefore, based on the type of gas-turbine unit, several different fuel lines of operation gas-turbine unit be can exist for
Road.Combustion system 200 disclosed in Fig. 2 and 3 is multistage premix combustion systems comprising the level Four of based on engine load is fired
Material injection.However, it is contemplated that going out, special fuel route and associated control mechanism be can be changed into including less or additional
Fuel circuit.
Referring now to Fig. 6-11, the additional detail in terms of combustion liner entrance area is drawn and discussed.First
Fig. 6 is turned to, the detailed view of the arrival end of the combustion liner of the prior art is shown.More specifically, combustion liner 600 has
The substantially ring-like ontology 602 of tape thickness 604, and the thermal boundary of the application of inner surface 608 along substantially ring-like ontology 602
Coating 606.Combustion liner 600 has arrival end 610.In the embodiment of the prior art, thermal barrier coating 606 extends to entrance
End 610, and it is formed together blunt nosed face 612.That is, arrival end 610 depends on being used for combustion liner for the embodiment of the prior art
600 sheet-metal thickness and there is up to 0.090 inch or bigger of combination thickness (metal+thermal barrier coating).When this burning
Bushing 600 together with Fig. 2-5 combustion system in use, combustion liner 600 and its arrival end 610 form blunt body, can fire
Unexpected result is generated when the stream of material and air is along and around the transmission of arrival end 610.More specifically, when fuel and air
When stream is transmitted around arrival end 610, the mixture of fuel and air is when it enters combustion liner 600 due to blunt body geometric form
Shape and tend to separate out.As understood by those skilled in the art, all flow separations so can help to for flame being anchored
At or near arrival end 610.The unexpected result causes the arrival end 610 of combustion liner 600 by the recirculation regions
The flame of formation corrodes, and causes the premature repairs or replacement to combustion liner.
The improvement to the arrival end 610 of the combustion liner of the prior art is depicted in Fig. 7.In an embodiment of the present invention,
Combustion liner 700 is set as the ontology 702 for having substantially ring-like, has the thickness T changed towards front area 704.Burning lining
Set 700 also has arrival end 706 and opposite outlet end (not shown).Substantially ring-like ontology 702 also has inner surface 708
With opposite outer surface, outer surface has the profile for determining shape, including the first outer surface 710 and second near arrival end 706
Outer surface 712, wherein the first outer surface 710 is located at the radial outside of the second outer surface 712.
The front area 704 of combustion liner 700 also has to fall from the first outer surface 710 towards arrival end 706 extends first
Angle 714, to reduce the thickness in front area 704 of combustion liner 700.For embodiment shown in fig. 7, first
Chamfering 714 is oriented with the angle of approximate 5-75 degree, and the thickness of combustion liner 700 is reduced to from approximate 0.1-0.25 inches into
Approximate 0.005-0.1 inches at mouth end 706.The rate of change of the thickness of chamfer angle, resulting thickness and combustion liner
It is only representative, and be not intended to limit the scope of the invention.As will be appreciated by one skilled in the art, combustion liner
Thickness, chamfer angle and the thickness change rate-compatible towards arrival end 706.However, by making via the first chamfering 714
Thickness change is tapered with first rate, the stream of the more fuel and air transmitted along the outer surface of substantially ring-like ontology 702
Design with the prior art is remained attached on the contrary on ring body 702.
Combustion liner 700 further includes the coating 716 being applied on the inner surface 708 of substantially ring-like ontology 702.For
A kind of this type coating of combustion liner 700 is thermal barrier coating.The thermal barrier coating 716 being applied on inner surface 708 includes that bonding applies
Layer 718 and ceramic Topcoating 720.For example, bonding coat 718 can apply approximate 0.001-0.010 inch thickness, and ceramics push up
Portion's coating 720 can apply approximation 0.010-0.200 inch thickness on bonding coat 718.As it will be understood by those skilled in the art that
As, thermal barrier coating can be normal business coating discussed above, or can also be more advanced thermal barrier coating, such as, fine and close
Vertical cracking of coating.As seen from Fig. 7, a part near arrival end 706 of coating is via with the angle of 5-75 degree
Orientation the second chamfering 722 and it is tapered, this keeps coating layer thickness tapered with the second rate towards arrival end 706.Second chamfering 722 can be through
It is formed by machining process, such as, is ground to the coating applied before or it can be due to applying the bonding coat applied and thermal boundary
The layer of layer is tapered and is formed.
Therefore, as seen from Figure 7, the first chamfering 714 and the second chamfering 722 form the blunt nosed of reduction at arrival end 706
Body region 724.In an embodiment of the present invention, the blunt nosed body region 724 of reduction has approximate 0.020 inch of thickness.However,
The expectation that other blunt nosed body regions 724 reduced may depend on combustion liner 700 constructs to use.As discussed before,
Blunt nosed body region generates recirculation regions.However, chamfer angle 714 and 722 of the invention reduces the size in this region, so as to
Reduce the trend that the stream of fuel and air is separated when it is transmitted towards arrival end 706.
However, using the blunt nosed body region 724 of the reduction formed by the present invention, along the outer of substantially ring-like ontology 702
The stream of the fuel and air of portion region transmission is kept along tapered surface 714 and 722, thus reduce the blunt body of the prior art
Unfavorable effect.
In alternative embodiment of the invention, the chamfering at bushing inlet end 706 alternatively includes such as institute in Fig. 8-10
The part of the rounding at the bushing inlet end shown or the blunt nosed body region of rounding.More specifically, and as illustrated in figs. 8-10, burn
Bushing 800 has arrival end 806, and substitutes chamfer angle 714 and 722 shown in fig. 7, and combustion liner 800 is in arrival end 806
Place has one or more radiuses.That is, combustion liner 800 includes having arrival end 806 and outlet end (not shown) generally
The ontology 802 of annular.Ring body 802 has inner surface 808 and outer surface 810.In this embodiment, inner surface 808 has
It is applied to thermal barrier coating 820 thereon.However, the embodiment, which is included at bushing inlet, to be formed to combustion unlike the embodiment of Fig. 7
Burn one or more radiuses in bushing 800.More specifically, in fig. 8, which includes substantially ring-like
Ontology 802 near arrival end 806 surround outer surface 810 radius R.Radius R may depend on many factors and change.So
And, it is preferred that radius R extends certain distance to extend the length of the tapered surface 714 of the embodiment in essentially equal to Fig. 7
Degree.Therefore, radius R covers the same general areas of tapered surface 714.However, although radius provides and 714 class of tapered surface
As benefit, but it is advantageous not as good as tapered surface 714.Radius R mentions the risk of air flow separation due to curved surface
It is high.It is kept in this region in addition, this radius negatively affects any flame.
Alternatively, and as shown in Figure 9, the one or more radius R of combustion liner 800 can at arrival end 806 edge
The formation of thermal barrier coating 820 being applied on inner surface 808.The radius R of thermal barrier coating 820 may depend on coating layer thickness and change.
Such as the embodiment of Fig. 8, the radius R of thermal barrier coating also impacts negatively on flame and is maintained in arrival end 806.
Then, referring to Fig.1 0, one or more radius R may include the first radius R1 and the second radius R2.It is more specific and
Speech, substantially ring-like ontology 802 have the first radius R1 for the radius R2 for being substantially greater than thermal barrier coating 820.Therefore, entrance
Hold collectively forming and the bull nose of the inlet of combustion liner (bullnose) comparable shape for the R1 and R2 at 806.
The Blunt leading edge of combustion liner needed for construction disclosed in Fig. 8-10 provides bush structure integrality.However,
Reduce leading edge thickness by reducing the too early thermal wear for keeping the trend of flame to prevent combustion liner arrival end 806.Radius
R, R1 and/or R2 is preferably formed by the process of lapping to bushing and/or thermal barrier coating.
Referring now to fig. 11, disclose the method 1100 for reducing the recirculation regions in gas turbine combustor.More specifically
For, in step 1102, combustion liner is set as having the chamfering along the outer surface of combustion liner, be applied in combustion liner
The chamfering of the coating on coating and inner surface on surface.Then, in step 1104, the mixture of fuel and air
It is guided along the outer surface of combustion liner.Then the mixture of fuel and air surrounds the arrival end of combustion liner in a step 1106
It turns to, so that mixture is kept at least close to the chamfered part of combustion liner.Then, in step 1108, fuel and air
Mixture is directed in combustion liner, at this, is lighted and is supplied power to gas-turbine unit.
Although describing the present invention in the embodiment for being known as currently preferred embodiment, it is to be understood that of the invention
It is not limited to disclosed embodiment, and it is opposite, it is intended that various modifications and equivalent arrangements in covering the scope of the following claims.It closes
Describe the present invention in specific embodiment, be intended to all be in all respects it is exemplary and not restrictive.
It will be seen that the present invention is adapted for reaching together with the upper of other clear in system and method and intrinsic advantages from front
The invention of all purposes and target that text illustrates.It will be appreciated that certain features and sub-portfolio are practical, and can without reference to
It is used in the case where other features and sub-portfolio.This is envisioned and in the range by the scope of the claims.
Claims (21)
1. a kind of combustion liner, comprising: substantially ring-like ontology, with thickness, arrival end and opposite outlet end, institute
Substantially ring-like ontology is stated with inner surface and opposite outer surface, the outer surface has surely outer near the arrival end
The profile of shape, so that the outer surface includes the first outer surface and the second outer surface, wherein first outer surface is positioned at described
The radial outside of second outer surface, and the first chamfering extends to the arrival end from first outer surface；And it is applied to institute
State the coating on inner surface, wherein the coating includes bonding coat and ceramic Topcoating, the coating in the entrance
At least part near end has the second chamfering, to keep coating layer thickness tapered towards the arrival end.
2. combustion liner according to claim 1, which is characterized in that the combustion liner is in first outer surface
Thickness is greater than the combustion liner in the thickness of second outer surface.
3. combustion liner according to claim 1, which is characterized in that first chamfering and second chamfering are described
The blunt nosed body region reduced is formed at arrival end.
4. combustion liner according to claim 3, which is characterized in that the blunt body at the arrival end is served as a contrast in the burning
The arrival end of set is provided about recirculation regions.
5. combustion liner according to claim 1, which is characterized in that first chamfering, which has, is oriented approximate 5-75 degree
Chamfer angle.
6. combustion liner according to claim 1, which is characterized in that the institute being applied on the inner surface of the combustion liner
Stating coating has fine and close vertical cracking of micro-structure.
7. a kind of intake section of combustion liner, comprising: the substantially ring-like ontology tapered from the first insert thickness has
Second insert thickness, and it is tapered with first rate towards arrival end from the first insert thickness；And it is applied to described substantially ring-like
Ontology inner wall on coating, at least part of the coating is tapered to from the arrival end from first coating thickness
Two coating layer thicknesses, the coating are tapered with the second rate.
8. intake section according to claim 7, which is characterized in that the arrival end has approximation 0.005-0.100 English
Very little thickness.
9. intake section according to claim 7, which is characterized in that the thickness of combustion liner is with the angle of approximate 5-75 degree
It is tapered.
10. intake section according to claim 7, which is characterized in that the thickness of the coating is with the angle of approximate 5-75 degree
It spends tapered.
11. intake section according to claim 7, which is characterized in that the coating has the vertical cracking of micro- of densification
Structure.
12. intake section according to claim 7, which is characterized in that the first coating is with a thickness of 0.010 inch approximate
To 0.200 inch.
13. a kind of method for reducing the recirculation regions in combustion liner, comprising: provide combustion liner, have along the combustion
The chamfering for burning the outer surface of bushing, the coating that is applied on the inner surface of the combustion liner and on the inner surface
At least part of chamfering of coating；Along the mixture of the outer surface of combustion liner guidance fuel and air；Make the combustion
The mixture of material and air is turned to around the arrival end of the combustion liner, so that the mixture is kept at least close to the combustion
Burn the chamfered part of bushing；And the mixture is directed in the combustion liner.
14. according to the method for claim 13, which is characterized in that the arrival end forms blunt body, compared to described
The thickness of combustion liner and the coating has reduced thickness.
15. according to the method for claim 14, which is characterized in that the blunt body has 0.005-0.050 inches of approximation
Thickness.
16. a kind of combustion liner, comprising: substantially ring-like ontology, with thickness, arrival end and opposite outlet end, institute
Substantially ring-like ontology is stated with inner surface and opposite outer surface, shape is determined according to the first radius in the outer surface；And
The coating being applied on the inner surface, wherein the coating includes bonding coat and ceramic Topcoating, the coating is extremely
Few a part determines shape according to the second radius, so that first radius incorporates second radius at the arrival end.
17. combustion liner according to claim 16, which is characterized in that first radius and the second radius phase
It cuts.
18. combustion liner according to claim 16, which is characterized in that first radius is greater than second radius.
19. combustion liner according to claim 16, which is characterized in that second radius is greater than first radius.
20. a kind of combustion liner, comprising: substantially ring-like ontology, with thickness, arrival end and opposite outlet end, institute
Substantially ring-like ontology is stated with inner surface and opposite outer surface, the outer surface has surely outer near the arrival end
The profile of shape, so that the outer surface includes the first outer surface and the second outer surface, wherein first outer surface is positioned at described
The radial outside of second outer surface, and the first chamfering extends to the arrival end from first outer surface；And it is applied to institute
State the coating on inner surface, wherein the coating includes bonding coat and ceramic Topcoating, the coating in the entrance
At least part near end has the radius at the arrival end.
21. a kind of combustion liner, comprising: substantially ring-like ontology, with thickness, arrival end and opposite outlet end, institute
Substantially ring-like ontology is stated with inner surface and opposite outer surface, shape is determined according to radius in the outer surface；And apply
Coating onto the inner surface, wherein the coating includes bonding coat and ceramic Topcoating, the coating described
At least part near arrival end has chamfering, to keep coating layer thickness tapered towards the arrival end.
Applications Claiming Priority (3)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US14/549922 | 2014-11-21 | ||
| 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 (2)
| Publication Number | Publication Date |
|---|---|
| CN107429920A CN107429920A (en) | 2017-12-01 |
| CN107429920B true CN107429920B (en) | 2019-11-05 |
Family
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Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201580074243.2A Active CN107429920B (en) | 2014-11-21 | 2015-11-20 | Flame front burner determines the bushing of shape |
Country Status (5)
| Country | Link |
|---|---|
| EP (1) | EP3221643B1 (en) |
| JP (1) | JP6824165B2 (en) |
| KR (1) | KR102405991B1 (en) |
| CN (1) | CN107429920B (en) |
| WO (1) | WO2016099805A2 (en) |
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| 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 |
| US11859819B2 (en) | 2021-10-15 | 2024-01-02 | General Electric Company | Ceramic composite combustor dome and liners |
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Also Published As
| Publication number | Publication date |
|---|---|
| JP6824165B2 (en) | 2021-02-03 |
| JP2017536524A (en) | 2017-12-07 |
| KR102405991B1 (en) | 2022-06-10 |
| WO2016099805A3 (en) | 2016-10-27 |
| EP3221643A4 (en) | 2018-09-05 |
| EP3221643A2 (en) | 2017-09-27 |
| EP3221643B1 (en) | 2020-02-26 |
| WO2016099805A2 (en) | 2016-06-23 |
| KR20170106302A (en) | 2017-09-20 |
| CN107429920A (en) | 2017-12-01 |
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