CN103998745B - Can-annular combustor with staged and tangential fuel-air nozzles for use on gas turbine engines - Google Patents
Can-annular combustor with staged and tangential fuel-air nozzles for use on gas turbine engines Download PDFInfo
- Publication number
- CN103998745B CN103998745B CN201180073013.6A CN201180073013A CN103998745B CN 103998745 B CN103998745 B CN 103998745B CN 201180073013 A CN201180073013 A CN 201180073013A CN 103998745 B CN103998745 B CN 103998745B
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- Prior art keywords
- air
- nozzle
- jet
- fuel
- cylinder volume
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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
- F23R3/28—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
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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
- F23R3/44—Combustion chambers comprising a single tubular flame tube within a tubular casing
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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
- F23R3/58—Cyclone or vortex type combustion chambers
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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
- F23R3/50—Combustion chambers comprising an annular flame tube within an annular casing
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- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (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 fuel and air. Fuel, air and/or 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 an optimal 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
Technical field
The present invention relates to the equipment in gas turbine, described equipment contributes to accommodating and produces fuel and air mixture
Burning.This equipment is including but not limited to used in the military related application with commercial aircraft, generating and other gas turbines
Fuel-air nozzle, burner inner liner and housing and flowing transitioning component.
Background technology
Gas turbine includes drawing the machinery of energy from the burning gases flowing with very high temperature, pressure and speed
Device.The energy being drawn can be used for driving the electromotor for generating electricity, or provides required thrust for aircraft.Typical combustion gas
Turbine is made up of compound compressor, and wherein air is compressed to high pressure.Then the air compressing is fired with specific in a combustion chamber
The mixing of empty ratio, its temperature of here increases.Then the burning gases of high temperature and high pressure expand through turbine to extract energy, with according to
There is provided required thrust by this application or drive electromotor.This turbine includes at least one single level, is banked leaf by one for every grade
Constitute with row's blade.This blade is circumferentially distributed on rotary hub, and wherein the height of each blade covers hot gas flow path.
Every grade of non-rotating blade circumferentially, and it also extends across hot gas flow path.The invention being comprised includes gas turbine
Combustor and assembly fuel and air being incorporated in described device.
The combustor section of gas turbine can have several different types：Tubular/tubulose, annular and above the two
The tubular toroidal combustion chamber combining to form.In the assembly, the fuel-air mixture of compression passes through Fuel-air cyclone
Or nozzle, and there is the combustion reaction of mixture, thus producing the thermal current leading to mixture density to decline, and in downstream plus
Speed.Tubular type combustor generally includes independent, multiple circumferentially-spaced cylinder, and described cylinder accommodates the flame of each nozzle respectively.
Subsequently, come from the air-flow of each and be conducted through conduit, and before it enters the first grade blade in annular transitional part
Integrate.In toroidal combustion chamber type, fuel-air nozzle is generally circumferentially distributed, and introduces a mixture into generation burning
In single toroidal combustion chamber.Air-flow only exits into the first stage turbine from the downstream of annuluss, without transition
Part.Key difference is that of last type, i.e. tubular toroidal combustion chamber, it has cylinder that is independent, being surrounded by toroidal shell,
Described toroidal shell accommodates the air being fed in each.According to the present invention, each deformation have it favourable and unfavorable it
Place.
In the combustor for gas turbine, due to many reasons, generally introducing will be vortexed by Fuel-air nozzle
To mixture.One reason is to improve mixability thus promoting to burn, Another reason is to increase vortex and has consolidated fire
Flame, to prevent fray-out of flame, and allows fuel-air mixture more thin, to reduce discharge.Fuel-air nozzle can be adopted
With different configurations, such as one, to multiple annular entries, each annular entrance has swirl vane.
For other gas turbine component, need to implement cooling means, to prevent combustor material from melting.A kind of cooling
The typical method of combustor is effusion cooling, surrounds combustion liner with extra offset liner and implements, wherein additionally
Offset liner be located between combustor and combustion liner, compressor bleed air is passed through blending hole and cooling duct and passes through to go forward side by side
Enter to hot gas flow path.This technology removes heat from assembly, forms cold air border between bushing and burning gases simultaneously
Layer film, thus stop the heat transfer to bushing.According to it, the axial location on bushing has two purposes to blending hole：Closer to
The blending hole of Fuel-air nozzle will be helpful to the mixing of gas, to promote to burn, provides unburned air for burning simultaneously,
Second, it is closer to the hole meeting cooling hot gas stream of turbine placement, and may be designed as controlling the outlet temperature of combustor to divide
Cloth.
As can be seen that different methods and techniques are incorporated in the design for the combustor of gas turbine, to promote
Burn and reduce discharge.As long as combustion gas turbine tends to producing less pollution than other electricity-generating methods, just also have in this area
Room for improvement.With the severization of the discharge standard of government in some countries, this technology needs to improve to meet these requirements.
Content of the invention
With regard to the present invention, it provides a kind of improved Combustion chamber design, and it can operate in the usual manner, make simultaneously
The burning produced pollution discharge of fuel and air mixture minimizes, and solve that this kind equipment faced some other
Problem.The present invention includes typical tubular toroidal combustion chamber, and it has fuel and air gas jets and/or blending hole, and fuel is empty
Compressor bleed air and pressurized fuel are incorporated into by the gas jets and/or blending hole diverse location in longitudinal direction and circumferential direction
Combustor.The original creation feature of the present invention is fuel and air nozzle is placed in such a manner：Produce and improve combustion reactant and product
The environment of the mixing of thing.By fuel and air nozzle classification, make the main spray fuel of upstream nozzle, another group of downstream nozzle is main
Injection air, this facilitates the mixing of combustion reactant, and combustion zone produce can greatly reduce formed NOx specific oxygen dense
Degree.In the device, not additional/fixing flame, but there is the region of adjacent front wall in cylinder, send out in this region
Raw diffusion combustion.The construction of detached fuel and air nozzle means that downstream injection and the air of upstream diffusion can be diluted,
Therefore reduce the oxygen concentration producing flame, and reduce flame peak temperature.This allows the invention to reduce discharge.In addition,
Combustion zone downstream introduces compressor bleed air so that all of CO being formed during burning enters first order turbine at it
Burned before machine/to consume.In fact, combustor can improve the emission level of combustion gas turbine, therefore reduce emission control and set
Standby needs, reduce the environmental effect of such equipment simultaneously.Except these improvement, in each cylinder, tangential firing fuel and combustion
Initial flame front is directed to adjacent burner nozzle, thus substantially increasing the igniting of combustor by material air nozzle
Journey.
Brief description
Referring to accompanying drawing：
Fig. 1 is the two-dimentional sketch of the tubular ring device illustrating to have the nozzle being attached to outside barrel liner, this nozzle
Fuel and air are ejected in common plane；
Fig. 2 is the two-dimentional sketch illustrating to be applied to the basic conception of tangential nozzle of cylinder in tubular toroidal combustion chamber；
Fig. 3 is the isometric side view of the upstream portion of the exemplary configuration of described invention；
Fig. 4 A is the isometric cross-sectional view of the present invention；
Fig. 4 B is the close up view of Fig. 4 A image；
Fig. 5 is the profile of part A-A illustrating to limit in Fig. 3；And
Fig. 6 is the profile of part B-B illustrating to limit in Fig. 3.
Specific embodiment
Fig. 1 illustrates the embodiment of the usual configuration of tubular toroidal combustion chamber, and it has circumferentially spaced in common radius
Cylinder 1, all of cylinder is enclosed between tubular outer lining 2 and tubular liner 3.Fig. 1 also shows the tangential nozzle of cylinder
Arrangement.Fig. 2 illustrates the more details of cylinder.Barrel liner 4 forms cylinder volume, fuel/air mixture nozzle 5 spray fuel or sky
Gas.Nozzle angulation 8 between the tangent line of nozzle centerline 6 and the intersection of barrel liner 4 and nozzle centerline 6.
This angle defines the circumference of nozzle.
Fig. 2 also show the usual operation of the example middle cylinder body of tubular annular firing chamber device, wherein fuel or air 9 with
Angle 8 sprays into cylinder 1.In the present invention and unfixed flame 10 is formed in the path 11 consistent with barrel liner, and wear
Cross cylinder.These tangential nozzles make the air-flow of each nozzle and adjacent downstream nozzle interact.This key features is led to
Cross fuel at the nozzle of flame ignition adjacent downstream of a nozzle for the permission to promote to light a fire, and reduce for multiple combustions
The problem of burner nozzle orientation.
Fig. 3 shows the initial or upstream portion of the cylinder of example removing downstream part.The present invention has along cylinder
Longitudinally spaced multiple nozzles rows.Often row's nozzle 12,13 can have at least one nozzle, and with respect to adjacent nozzle row
The circumferential angle of skew one.Especially, the nozzle 12 in the nozzle row of antetheca 15 is in the foregoing manner by pure fuel/major part fuel
It is ejected in cylinder, pure compression machine is discharged air or fuel-air by the nozzle 13 in these nozzle downstreams in a similar manner simultaneously
Mixture is ejected in cylinder.This cylinder can also have hole 14 or the passage of multiple rows of circumferentially spaced, so that cooling air is in office
One position enters cylinder.
Fig. 4 A and Fig. 4 B illustrates the most upstream face 15 of cylinder, and it can have the hole 16 similar with blending hole, and described hole 16 permits
Permitted compressor and exhausted air to cylinder.Fig. 5 and Fig. 6 illustrate nozzle 12,13 in every group of nozzle as how necessarily circumferential angle inclined
Put.The nozzle of different rows allows the sequential fuel injection of fuel and air, produces the burning occurring without high oxygen concentration in antetheca vicinity
Reactant area, this will reduce flame peak temperature.To dilute from combustion product towards the flue gas that antetheca is upstream advanced so that
Combustion reactant can meet relatively low oxygen concentration.Can be reduced by this burning situation that classification fuel and air nozzle produce
Discharge.
The present invention is described above according to preferred embodiment.However, one skilled in the art will recognize that without departing from this
Embodiment described above can be made when inventive nature and scope with change and modification.Here for illustration purposes right selecting
The different change of embodiment is easy to be expected by those skilled in the art with modification.To a certain extent, such modification and variant be simultaneously
Without departing from the spirit of the present invention, they also aim to be included within the scope of the present invention.
The present invention is adequately described with clear, specific nomenclature, so that skilled artisans appreciate that
And implement this invention, the claim of the present invention is as follows.
Claims (7)
1. a kind of tubular toroidal combustion chamber, sets for continental rise generating equipment, the ground- or sea-based vehicles or aircraft engine
Gas turbine used in standby, including：The cylinder of multiple circumferentially spaceds, described cylinder is enclosed between two cylindrical sleeve,
Described cylinder limits single combustion zone and each cylinder is barrel liner, and described barrel liner has upstream end and downstream
End, described upstream end includes antetheca, and described combustion zone is the cylinder volume of described barrel liner, and described cylinder volume is in longitudinal direction
On direction, the described antetheca from the described upstream end of described barrel liner extends to the described downstream of described barrel liner；Pass through
Multiple blending hole of described antetheca, the plurality of blending hole allows compressor air-out in the longitudinal direction of described cylinder volume
On enter in described cylinder volume；First group of tangential sensing and the first jet of circumferentially spaced, described first jet is located at institute
State between the described upstream end of barrel liner and described downstream, one of composition of air and Fuel-air composition are sprayed
It is mapped in described cylinder volume；And the second nozzle of second group of tangential sensing and circumferentially spaced, described second nozzle is positioned at institute
State between first jet and the described upstream end of described barrel liner, with the institute in the described upstream end through described barrel liner
State and between the plurality of blending hole of antetheca and described first jet, propellant composition is ejected in described cylinder volume.
2. tubular toroidal combustion chamber according to claim 1, further includes the circumferentially spaced through described barrel liner
Cooling air hole, the cooling air hole of described circumferentially spaced is located at the described downstream of described barrel liner and described first spray
Between mouth, circumferentially to allow cooling air to enter institute between the described downstream of described cylinder volume and described first jet
State cylinder volume.
3. tubular toroidal combustion chamber according to claim 1, described first jet and described second nozzle do not extend to institute
State in cylinder volume.
4. tubular toroidal combustion chamber according to claim 1, wherein, described first jet guides any flame to next
Adjacent first jet the igniting contributing to each other, and described second nozzle any flame is guided adjacent to next
Second nozzle and the igniting that contributes to each other.
5. tubular toroidal combustion chamber according to claim 1, wherein, described first jet and described second nozzle promote
The mixing of the combusting component in described cylinder volume.
6. tubular toroidal combustion chamber according to claim 1, wherein, the plurality of blending hole allows compressor air-out
With less than described composition of air and the described Fuel-air composition being entered by first jet each described in described cylinder volume
In one velocity amplitude velocity amplitude pass through described barrel liner.
7. tubular toroidal combustion chamber according to claim 1, wherein, described first jet is with respect to described second nozzle week
To ground skew.
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| PCT/US2011/048612 WO2013028167A2 (en) | 2011-08-22 | 2011-08-22 | Can-annular combustor with staged and tangential fuel-air nozzles for use on gas turbine engines |
Publications (2)
| Publication Number | Publication Date |
|---|---|
| CN103998745A CN103998745A (en) | 2014-08-20 |
| CN103998745B true CN103998745B (en) | 2017-02-15 |
Family
ID=47747020
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN201180073013.6A Active CN103998745B (en) | 2011-08-22 | 2011-08-22 | Can-annular combustor with staged and tangential fuel-air nozzles for use on gas turbine engines |
Country Status (7)
| Country | Link |
|---|---|
| EP (1) | EP2748444B1 (en) |
| JP (1) | JP6086391B2 (en) |
| KR (1) | KR101774093B1 (en) |
| CN (1) | CN103998745B (en) |
| PL (1) | PL2748444T3 (en) |
| RU (1) | RU2611217C2 (en) |
| WO (1) | WO2013028167A2 (en) |
Families Citing this family (11)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US10139111B2 (en) * | 2014-03-28 | 2018-11-27 | Siemens Energy, Inc. | Dual outlet nozzle for a secondary fuel stage of a combustor of a gas turbine engine |
| FR3032781B1 (en) * | 2015-02-17 | 2018-07-06 | Safran Helicopter Engines | CONSTANT VOLUME COMBUSTION SYSTEM FOR AIRCRAFT ENGINE TURBOMACHINE |
| WO2018090383A1 (en) * | 2016-11-21 | 2018-05-24 | 深圳智慧能源技术有限公司 | Combustion chamber of gas turbine engine, and nozzle thereof |
| WO2018090384A1 (en) * | 2016-11-21 | 2018-05-24 | 深圳智慧能源技术有限公司 | Combustion chamber of gas turbine engine |
| CN106439914A (en) * | 2016-11-21 | 2017-02-22 | 深圳智慧能源技术有限公司 | Combustion chamber of combustion gas turbine |
| US11174792B2 (en) | 2019-05-21 | 2021-11-16 | General Electric Company | System and method for high frequency acoustic dampers with baffles |
| US11156164B2 (en) | 2019-05-21 | 2021-10-26 | General Electric Company | System and method for high frequency accoustic dampers with caps |
| KR102265626B1 (en) * | 2020-09-25 | 2021-06-16 | 박재현 | Sand spray test apparatus |
| CN114135901A (en) * | 2021-11-08 | 2022-03-04 | 中国航发四川燃气涡轮研究院 | Ablation-proof flame tube large-hole jet sleeve |
| CN114427689A (en) * | 2022-01-20 | 2022-05-03 | 中国空气动力研究与发展中心空天技术研究所 | Disc-shaped rotary detonation combustion chamber capable of realizing supersonic flow field observation |
| CN114857617B (en) * | 2022-05-20 | 2023-07-14 | 南昌航空大学 | Support plate flame stabilizer of band saw tooth type groove vortex generator |
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| US5113647A (en) * | 1989-12-22 | 1992-05-19 | Sundstrand Corporation | Gas turbine annular combustor |
| GB2295887A (en) * | 1994-12-08 | 1996-06-12 | Rolls Royce Plc | Combustor assembly |
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| JP4608154B2 (en) * | 2001-09-27 | 2011-01-05 | 大阪瓦斯株式会社 | Gas turbine combustion apparatus and gas turbine provided with the same |
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2011
- 2011-08-22 CN CN201180073013.6A patent/CN103998745B/en active Active
- 2011-08-22 EP EP11871243.9A patent/EP2748444B1/en active Active
- 2011-08-22 KR KR1020147007518A patent/KR101774093B1/en active IP Right Grant
- 2011-08-22 JP JP2014527126A patent/JP6086391B2/en active Active
- 2011-08-22 PL PL11871243T patent/PL2748444T3/en unknown
- 2011-08-22 WO PCT/US2011/048612 patent/WO2013028167A2/en active Search and Examination
- 2011-08-22 RU RU2014110628A patent/RU2611217C2/en active
Also Published As
| Publication number | Publication date |
|---|---|
| JP2014526029A (en) | 2014-10-02 |
| PL2748444T3 (en) | 2019-11-29 |
| KR20140082658A (en) | 2014-07-02 |
| JP6086391B2 (en) | 2017-03-01 |
| RU2014110628A (en) | 2015-09-27 |
| WO2013028167A2 (en) | 2013-02-28 |
| EP2748444A4 (en) | 2015-05-27 |
| EP2748444A2 (en) | 2014-07-02 |
| KR101774093B1 (en) | 2017-09-12 |
| RU2611217C2 (en) | 2017-02-21 |
| WO2013028167A3 (en) | 2014-03-20 |
| EP2748444B1 (en) | 2019-02-13 |
| CN103998745A (en) | 2014-08-20 |
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