U.S. patent application number 15/202751 was filed with the patent office on 2018-01-11 for deflector for gas turbine engine combustors and method of using the same.
The applicant listed for this patent is General Electric Company. Invention is credited to Manampathy Gangadharan Giridharan, Hiranya Kumar Nath.
Application Number | 20180010795 15/202751 |
Document ID | / |
Family ID | 60893225 |
Filed Date | 2018-01-11 |
United States Patent
Application |
20180010795 |
Kind Code |
A1 |
Nath; Hiranya Kumar ; et
al. |
January 11, 2018 |
DEFLECTOR FOR GAS TURBINE ENGINE COMBUSTORS AND METHOD OF USING THE
SAME
Abstract
A deflector for a gas turbine engine combustor. The combustor
includes a liner defining a combustion zone and a mixer assembly
configured to supply the combustion zone with a predetermined
mixture of fuel and air. The deflector includes a deflector body
configured to couple to the liner. The deflector body includes a
first surface configured to reflect thermal radiation to a
predetermined focal area, and an aperture extending through the
deflector body and configured to receive the mixer assembly
therethrough.
Inventors: |
Nath; Hiranya Kumar;
(Bangalore, IN) ; Giridharan; Manampathy Gangadharan;
(Cincinnati, OH) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
General Electric Company |
Schenectady |
NY |
US |
|
|
Family ID: |
60893225 |
Appl. No.: |
15/202751 |
Filed: |
July 6, 2016 |
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F23R 3/50 20130101; F23R
3/18 20130101; F23R 3/286 20130101; F23R 3/002 20130101 |
International
Class: |
F23R 3/00 20060101
F23R003/00 |
Claims
1. A deflector for a gas turbine engine combustor, the combustor
includes a liner defining a combustion zone and a mixer assembly
configured to supply the combustion zone with a predetermined
mixture of fuel and air, said deflector comprising: a deflector
body configured to couple to the liner and comprising a first
surface configured to reflect thermal radiation to a predetermined
focal area, wherein said deflector body defines an aperture
extending therethrough, the aperture configured to receive the
mixer assembly therethrough.
2. The deflector in accordance with claim 1, wherein said first
surface comprises a parabolic shape.
3. The deflector in accordance with claim 1, wherein the
predetermined focal area is positioned within the combustion
zone.
4. The deflector in accordance with claim 1, wherein the
predetermined focal area is configured to heat relatively cool
areas of the combustion zone to stabilize a flame during lean
combustion within the combustor.
5. The deflector in accordance with claim 1 further comprising a
thermal barrier coating formed on at least a portion of said
deflector body.
6. The deflector in accordance with claim 5, wherein said thermal
barrier coating is configured to reflect thermal radiation of a
flame to the predetermined focal area.
7. The deflector in accordance with claim 1, wherein said deflector
body comprises a superalloy substrate configured to reduce thermal
exposure to combustor components upstream of said deflector
body.
8. A combustor for a gas turbine engine comprising: a liner
defining a combustion zone; a mixer assembly configured to supply
the combustion zone with a predetermined mixture of fuel and air;
and a deflector coupled to said liner and comprising: a deflector
body configured to couple to the liner and comprising a first
surface configured to reflect incident thermal radiation to a
predetermined focal area, wherein said deflector body defines an
aperture extending therethrough, the aperture configured to receive
said mixer assembly therethrough.
9. The combustor in accordance with claim 8, wherein said first
surface comprises a parabolic shape.
10. The combustor in accordance with claim 8, wherein said
predetermined focal area is positioned within the combustion
zone.
11. The combustor in accordance with claim 8, wherein the
predetermined focal area is configured to heat relatively cool
areas of the combustion zone to stabilize a combustion flame during
lean combustion within the combustor.
12. The combustor in accordance with claim 8 further comprising a
thermal barrier coating formed on at least a portion of said
deflector body.
13. The combustor in accordance with claim 12, wherein said thermal
barrier coating is configured to reflect thermal radiation of a
combustion flame to the predetermined focal area.
14. The combustor in accordance with claim 8, wherein said
deflector body comprises a superalloy substrate configured to
reduce thermal exposure to combustor components upstream of said
deflector.
15. A method for stabilizing a flame within a gas turbine engine
combustor, the combustor including a liner defining a combustion
zone, a mixer assembly, and a deflector coupled to the liner, the
deflector including a deflector body configured to couple to the
liner, the deflector body including a first surface and defining an
aperture extending therethrough, the aperture configured to receive
the mixer assembly therethrough, said method comprising: generating
the flame within the combustion zone using a predetermined mixture
of fuel and air supplied by the mixture assembly; and reflecting
thermal radiation of the flame to a predetermined focal area.
16. The method in accordance with claim 15, wherein the first
surface has a parabolic shape.
17. The method in accordance with claim 16 further comprising
reflecting thermal radiation into the predetermined focal area
using the parabolic shape of the first surface.
18. The method in accordance with claim 15 further comprising
forming the first surface to focus the thermal radiation to the
predetermined focal area positioned within the combustion zone.
19. The method in accordance with claim 15 further comprising
forming a thermal barrier coating on at least a portion of the
deflector body.
20. The method in accordance with claim 15 further comprising
forming the body from a superalloy substrate configured to reduce
thermal exposure to combustor components upstream of the deflector.
Description
BACKGROUND
[0001] The field of the disclosure relates generally to gas turbine
engines and, more particularly, to deflectors for use in gas
turbine engine combustors.
[0002] Combustors are used to ignite fuel and air mixtures in gas
turbine engines to generate high energy working gases. Known
combustors include an outer liner and an inner liner defining an
annular combustion chamber in which the fuel and air are mixed and
burned. A dome mounted at the upstream end of the combustion
chamber includes mixers for mixing the fuel and air. Ignitors
mounted downstream from the mixers ignite the mixture, and it burns
in the combustion chamber. At least some combustors further include
a deflector coupled to the dome and surrounding the mixer that
prevent hot combustion gases produced within the combustion chamber
from impinging directly upon the dome and upstream components.
[0003] Air pollution concerns have led to stricter combustion
emissions standards. These standards regulate the emission of
nitrogen oxides (NOx), as well as other types of exhaust emissions,
from the gas turbine engine. Generally, NOx is formed during the
combustion process due to high flame temperatures in the combustor.
At least some combustors reduce NOx by using a lean fuel and air
mixture that tends to lower the flame temperature. For lean
mixtures, the volume percent of air in the mixture may typically be
increased up to a limit, sometimes referred to as a blowout limit,
at which the air and fuel mixture can no longer maintain a flame.
As such, to reduce NOx emissions, combustors typically use a lean
fuel and air mixture as close to the blowout limit as possible.
Lean fuel and air mixtures, however, may increase combustion
dynamics and flame instability because of the low temperature
flame. Flame instability and combustion dynamics generally reduces
combustor and overall gas turbine engine efficiency and
durability.
BRIEF DESCRIPTION
[0004] In one embodiment, a deflector for a gas turbine engine
combustor is provided. The combustor includes a liner defining a
combustion zone and a mixer assembly that is configured to supply
the combustion zone with a predetermined mixture of fuel and air.
The deflector includes a deflector body that is configured to
couple to the liner and including a first surface that is
configured to reflect thermal radiation to a predetermined focal
area. The deflector body defines an aperture extending
therethrough. The aperture is configured to receive the mixer
assembly therethrough.
[0005] In another embodiment, a combustor for a gas turbine engine
is provided. The combustor includes a liner defining a combustion
zone. A mixer assembly is configured to supply the combustion zone
with a predetermined mixture of fuel and air. The combustor further
includes a deflector coupled to the liner. The deflector includes a
deflector body that is configured to couple to the liner and
including a first surface that is configured to reflect thermal
radiation to a predetermined focal area. The deflector body defines
an aperture extending therethrough. The aperture is configured to
receive the mixer assembly therethrough.
[0006] In a further embodiment, a method for stabilizing a flame
within a gas turbine engine combustor is provided. The combustor
includes a liner defining a combustion zone, a mixer assembly, and
a deflector coupled to the liner. The deflector includes a
deflector body that is configured to couple to the liner. The
deflector body includes a first surface and defines an aperture
extending therethrough. The aperture is configured to receive the
mixer assembly therethrough. The method includes generating the
flame within the combustion zone using a predetermined mixture of
fuel and air supplied by the mixture assembly. The method further
includes reflecting thermal radiation of the flame to a
predetermined focal area.
DRAWINGS
[0007] These and other features, aspects, and advantages of the
present disclosure will become better understood when the following
detailed description is read with reference to the accompanying
drawings in which like characters represent like parts throughout
the drawings, wherein:
[0008] FIG. 1 is a schematic, cross-sectional illustration of an
exemplary turbofan engine in accordance with an example embodiment
of the present disclosure.
[0009] FIG. 2 is a cross-sectional view of an exemplary combustor
that may be used with the turbofan engine shown in FIG. 1.
[0010] FIG. 3 is a perspective view of an exemplary deflector that
may be used with the combustor shown in FIG. 2.
[0011] Unless otherwise indicated, the drawings provided herein are
meant to illustrate features of embodiments of this disclosure.
These features are believed to be applicable in a wide variety of
systems comprising one or more embodiments of this disclosure. As
such, the drawings are not meant to include all conventional
features known by those of ordinary skill in the art to be required
for the practice of the embodiments disclosed herein.
DETAILED DESCRIPTION
[0012] In the following specification and claims, reference will be
made to a number of terms, which shall be defined to have the
following meanings.
[0013] The singular forms "a," "an," and "the" include plural
references unless the context clearly dictates otherwise.
[0014] "Optional" or "optionally" means that the subsequently
described event or circumstance may or may not occur, and that the
description includes instances where the event occurs and instances
where it does not.
[0015] Approximating language, as used herein throughout the
specification and claims, may be applied to modify any quantitative
representation that could permissibly vary without resulting in a
change in the basic function to which it is related. Accordingly, a
value modified by a term or terms, such as "about,"
"approximately," and "substantially," are not to be limited to the
precise value specified. In at least some instances, the
approximating language may correspond to the precision of an
instrument for measuring the value. Here and throughout the
specification and claims, range limitations may be combined and/or
interchanged; such ranges are identified and include all the
sub-ranges contained therein unless context or language indicates
otherwise.
[0016] Embodiments of a deflector for a gas turbine engine
combustor as described herein provide an apparatus for lean
combustion systems that facilitates increasing combustor flame
stability and reducing emissions thereof. Specifically, the
deflector includes a body with a downstream surface that reflects
thermal radiation of a combustor flame within the combustor to a
predetermined focal area. The downstream surface is formed in a
predetermined shape, such as but not limited to, a parabolic shape
to facilitate thermal radiation reflection from the downstream
surface to the predetermined focal area. Furthermore, the deflector
may also include a reflective thermal barrier coating that further
facilitates reflecting thermal radiation of the combustion flame.
By reflecting thermal radiation from the deflector and back into
the combustion flame, reactants and products within cooler areas of
the combustion flame may increase in temperature to facilitate
stabilizing the combustion flame while maintaining a lean fuel/air
mixture. Stabilizing the combustion flame reduces combustion
dynamics and allows for the combustor to operate closer to a lean
blowout limit to reduce NOx emissions.
[0017] FIG. 1 is a schematic cross-sectional view of a gas turbine
engine in accordance with an exemplary embodiment of the present
disclosure. In the exemplary embodiment, the gas turbine engine is
a high-bypass turbofan jet engine 110, referred to herein as
"turbofan engine 110." As shown in FIG. 1, turbofan engine 110
defines an axial direction A (extending parallel to a longitudinal
centerline 112 provided for reference) and a radial direction R
(extending perpendicular to longitudinal centerline 112). In
general, turbofan engine 110 includes a fan case assembly 114 and a
gas turbine engine 116 disposed downstream from fan case assembly
114.
[0018] Gas turbine engine 116 includes a substantially tubular
outer casing 118 that defines an annular inlet 120. Outer casing
118 encases, in serial flow relationship, a compressor section
including a booster or low pressure (LP) compressor 122 and a high
pressure (HP) compressor 124; an annular combustion section 126
including a plurality of circumferentially spaced fuel nozzle
assemblies 218 (shown in FIG. 2); a turbine section including a
high pressure (HP) turbine 128 and a low pressure (LP) turbine 130;
and a jet exhaust nozzle section 132. A high pressure (HP) shaft or
spool 134 drivingly connects HP turbine 128 to HP compressor 124. A
low pressure (LP) shaft or spool 136 drivingly connects LP turbine
130 to LP compressor 122. The compressor section, combustion
section 126, turbine section, and exhaust nozzle section 132
together define an air flow path 138.
[0019] In the exemplary embodiment, fan case assembly 114 includes
a fan 140 having a plurality of fan blades 142 coupled to a disk
144 in a spaced apart manner. As depicted, fan blades 142 extend
outwardly from disk 144 generally along radial direction R. Fan
blades 142 and disk 144 are together rotatable about longitudinal
centerline 112 by LP shaft 136.
[0020] Referring still to the exemplary embodiment of FIG. 1, disk
144 is covered by a rotatable front hub 146 aerodynamically
contoured to promote an airflow through plurality of fan blades
142. Additionally, exemplary fan case assembly 114 includes an
annular fan casing or outer nacelle 150 that circumferentially
surrounds fan 140 and/or at least a portion of gas turbine engine
116. Nacelle 150 is supported relative to gas turbine engine 116 by
an outlet guide vane (OGV) assembly 152. Moreover, a downstream
section 154 of nacelle 150 may extend over an outer portion of gas
turbine engine 116 so as to define a bypass airflow duct 156
between nacelle 150 and outer casing 118.
[0021] During operation of turbofan engine 110, a volume of air 158
enters turbofan engine 110 through an associated inlet 160 of
nacelle 150 and/or fan case assembly 114. As air 158 passes across
fan blades 142, a first portion of air 158 as indicated by arrow
162, known as fan stream air flow, is directed or routed into
bypass airflow duct 156 and a second portion of air 158 as
indicated by arrow 164 is directed or routed into air flow path
138, or more specifically into booster compressor 122. The ratio
between first portion of air 162 and second portion of air 164 is
commonly known as a bypass ratio. The pressure of second portion of
air 164 is then increased, forming compressed air 166, as it is
routed through booster compressor 122 and HP compressor 124 and
into combustion section 126, where it is mixed with fuel 168 and
burned to provide combustion gases 170.
[0022] Combustion gases 170 are routed through HP turbine 128 where
a portion of thermal and/or kinetic energy from combustion gases
170 is extracted via sequential stages of HP turbine stator vanes
172 that are coupled to outer casing 118 and HP turbine rotor
blades 174 that are coupled to HP shaft or spool 134, thus causing
HP shaft or spool 134 to rotate, thereby supporting operation of HP
compressor 124. Combustion gases 170 are then routed through LP
turbine 130 where a second portion of thermal and kinetic energy is
extracted from combustion gases 170 via sequential stages of LP
turbine stator vanes 176 that are coupled to outer casing 118 and
LP turbine rotor blades 178 that are coupled to LP shaft or spool
136, thus causing LP shaft or spool 136 to rotate, thereby
supporting operation of booster compressor 122 and/or rotation of
fan 140. Combustion gases 170 are subsequently routed through jet
exhaust nozzle section 132 of gas turbine engine 116 to provide
propulsive thrust. HP turbine 128, LP turbine 130, and jet exhaust
nozzle section 132 at least partially define a hot gas path 180 for
routing combustion gases 168 through gas turbine engine 116.
Simultaneously, the pressure of fan stream air 162 is substantially
increased as fan stream air 162 is routed through bypass airflow
duct 156, including through outlet guide vane assembly 152 before
it is exhausted from a fan nozzle exhaust section 182 of turbofan
engine 110, also providing propulsive thrust.
[0023] It should be appreciated, however, that the exemplary
turbofan engine 110 depicted in FIG. 1 is by way of example only,
and that in other exemplary embodiments, turbofan engine 110 may
have any other suitable configuration. It should also be
appreciated, that in still other exemplary embodiments, aspects of
the present disclosure may be incorporated into any other suitable
gas turbine engine. For example, in other exemplary embodiments,
aspects of the present disclosure may be incorporated into, e.g., a
turboprop engine, a military purpose engine, and a marine or
land-based aero-derivative engine.
[0024] FIG. 2 is a cross-sectional view of an exemplary combustor
200 that may be used with turbofan engine 110 (shown in FIG. 1). In
the exemplary embodiment, annular combustion section 126 includes a
combustor 200 having a combustion zone or chamber 202 defined by
annular, radially outer and radially inner liners 204 and 206.
Specifically, outer liner 204 defines a radially outer boundary of
combustion chamber 202, and inner liner 206 defines a radially
inner boundary of combustion chamber 202. Liners 204 and 206 are
spaced radially inward from an annular combustor casing 208 which
extends circumferentially around liners 204 and 206. Combustor 200
also includes an annular dome 210 mounted upstream from outer and
inner liners 204 and 206 respectively. Dome 210 defines an upstream
end of combustion chamber 202 and a plurality of mixer assemblies
212 are spaced circumferentially around dome 210 to deliver a
mixture of fuel and air to combustion chamber 202. Each mixer
assembly 212 includes a pilot mixer 214 and a main mixer 216. Main
mixer 216 is concentrically aligned with respect to pilot mixer 214
and extends circumferentially around pilot mixer 214. A plurality
of circumferentially spaced and axially-extending fuel nozzle
assemblies 218 are coupled in flow communication with each
respective mixer assembly 212. Furthermore, in the exemplary
embodiment, combustor 200 includes one or more deflectors 220 that
are coupled to and spaced circumferentially around dome 210 at each
mixer assembly 212 location. Downstream of mixer assembly 212 and
deflector 220 is an igniter 222 that extends through outer casing
208 and into combustion chamber 202 to provide initial ignition of
the mixture of compressed air 166 and fuel 168. In various
embodiments, igniter 222 may provide continuous or intermittent
ignition support to combustion chamber 202.
[0025] In operation, combustor 200 receives compressed air 166
discharged from HP compressor 124 in a diffuser section 224 at flow
upstream of combustion chamber 202. A portion of the flow of
compressed air 166 is channeled through mixer assembly 212. At
mixer assembly 212 compressed air 166 is mixed with fuel 168 from
fuel nozzle assembly 218 and discharged into combustion chamber 202
where the mixture of air 166 and fuel 168 is ignited by igniter 222
creating a flame 224 within combustion chamber 202 that burns the
mixture and provides combustion gases 170 that are channeled
downstream to HP turbine 128 (shown in FIG. 1). In the exemplary
embodiment, combustor 200 is a lean combustor. Specifically, at
engine start conditions and engine low power operation, combustor
200 uses only fuel 168 provided to the pilot mixer 214 for
generating combustion gases 170. At pilot mixer 214, fuel 168
includes a pilot fuel stream 226 that is mixed with a first portion
228 of compressed air 166 to provide a rich mixture (higher fuel
226 to air 228 ratios within the mixture) that is ignited for a
pilot flame 230 within a region 232 that is adjacent to pilot mixer
214. At engine high power operation, combustor 200 uses fuel 168
split between pilot mixer 214 and main mixer 216 for generating
combustion gases 170. At main mixer 216, fuel 168 includes a main
fuel stream 234 that is mixed with a second portion 236 of
compressed air 166 to provide a lean mixture (lower fuel 234 to air
236 ratios within the fuel-air mixture) that is ignited for a main
flame 238 within a region 240 that is adjacent to main mixer 216.
At engine high power operation most of fuel 168 is injected through
main mixer 216 thus providing a lean burn combustion process to
generate combustion gases 170 while reducing NOx emissions.
[0026] Generally, pilot flame 230 burns at a higher temperature
than main flame 238 because the fuel 226 air 228 mixtures are
richer. As such, during engine start and engine low power operation
combustion dynamics are low leading to pilot flame 230/flame 224
that is stable. Main flame 238, however, generally burns at a lower
temperature than pilot flame 230 because the fuel 234 air 236
mixtures are leaner. As such, during high power engine operation,
when most of fuel 168 is injected through main mixer 216, main
flame 238/flame 224 instability may occur due to lower temperatures
leading to combustion dynamics. To facilitate reducing NOx
emissions from the combustion process, flame 224 temperatures
during high power engine operation are reduced as low as possible
and close to a blowout limit by increasing the air 236 to fuel 234
ratio because NOx is formed at high flame temperatures. As such,
cooler temperature regions, such as region 242, are formed within
flame stabilization zones that decreases flame 224 stability and
increases a likelihood of flame blowout.
[0027] In the exemplary embodiment, deflector 220 facilitates
increasing flame 224 stability and decreases combustion dynamics
and a likelihood of flame blowout during engine operation at high
power levels while using the lean burn combustion process.
Specifically, deflector 220 reflects 244 thermal radiation and
infrared radiation generated by flame 224 within combustion chamber
202 to a predetermined focal area 246 within cooler temperature
region 242 to increase the temperature of region 242. Flame 224
generates thermal radiation by burning the fuel 168 and air 166
mixture which typically heats up the surrounding combustor
components, such as outer and inner liners 204 and 206 and dome
210. By reflecting 244 the thermal radiation back into flame 224,
the temperature of region 242 increases. The temperature of region
242 increases by heating up entrained carbon dioxide and water
vapor therein. Thus, reducing combustion dynamics and a likelihood
of flame blowout, and increasing flame 224 stability.
[0028] It should be appreciated, that the exemplary combustor 200
illustrated in FIG. 2 is by way of example only, and that in other
exemplary embodiments, combustor 200 may have any other suitable
configuration for lean based combustion. It should further be
appreciated, that in still other exemplary embodiments, aspects of
the present disclosure may be incorporated into any other suitable
gas turbine engine including land-based aero-derivative engine
combustion systems.
[0029] FIG. 3 is a perspective view of an exemplary deflector 220
that may be used with combustor 200 (shown in FIG. 2). In the
exemplary embodiment, deflector 220 is substantially circular and
includes a body 300 formed with an aperture 302 sized to at least
partially receive mixer assembly 212 (shown in FIG. 2). Body 300 is
coupled to dome 210 (shown in FIG. 2), for example, by brazing.
Body 300 includes an upstream surface 304 and a downstream surface
306. Downstream surface 306 facilitates reflecting thermal
radiation from combustor flame 224 (shown in FIG. 2) to
predetermined focal area 246 (shown in FIG. 2). For example,
downstream surface 306 is parabolic such that downstream surface
306 reflects thermal radiation from combustor flame 224 to
predetermined focal area 246.
[0030] Parabolic shape 308 of deflector 220 receives thermal
radiation on the curved downstream surface 306 and reflects the
thermal radiation to predetermined focal area 246. Parabolic
curvature 308 of downstream surface 306 may be sized to position
focal area 246 at any location that facilitates increasing
temperature of combustor flame 224. Thus, reducing combustion
dynamics and a likelihood of flame blowout, and increasing flame
224 stability. In alternative embodiments, downstream surface 306
may have any other shape/contour that enables deflector 220 to
functions as described herein. Additionally, in alternative
embodiments, the substantially circular deflector body 300 may be
trimmed to form a polygonal periphery.
[0031] In the exemplary embodiment, body 300 is fabricated from a
superalloy substrate 310 and coated with a thermal barrier coating
312 to reduce thermal exposure when combustor 200 is operating.
Physical vapor deposition thermal barrier coating, TBC 312, is
applied to deflector 220 and provides thermal protection thereto.
Furthermore, TBC 312 facilitates reflecting thermal radiation of
flame 224 as described above.
[0032] During operation of combustor 200, deflector 220 protects
dome 210 and mixer assembly 212 from hot gases and thermal flame
radiation generated with combustion chamber 202. Furthermore,
parabolic curvature 308 and TBC 312 reflect the thermal flame
radiation back into flame 224 to increase flame stability during
lean combustor operation.
[0033] The above-described embodiments of a deflector for a gas
turbine engine combustor provide an apparatus for lean combustion
systems that facilitates increasing combustor flame stability and
reducing emissions thereof. Specifically, the deflector includes a
body with a downstream surface that reflects thermal radiation of a
combustor flame within the combustor to a predetermined focal area.
The downstream surface is formed in a predetermined shape, such as,
but not limited to, a parabolic shape to facilitate thermal
radiation reflection from the downstream surface to the
predetermined focal area. Furthermore, the deflector may also
include a reflective thermal barrier coating that further
facilitates reflecting thermal radiation of the combustion flame.
By reflecting thermal radiation from the deflector and back into
the combustion flame, reactants and products within cooler areas of
the combustion flame may increase in temperature to facilitate
stabilizing the combustion flame while maintaining a lean fuel/air
mixture. Stabilizing the combustion flame reduces combustion
dynamics and allows for the combustor to operate closer to a lean
blowout limit to reduce NOx emissions.
[0034] An exemplary technical effect of the methods, systems, and
apparatus described herein includes at least one of: (a) reducing
combustion dynamics in a gas turbine combustion system; (b)
increasing lean combustion flame temperature; (c) increasing
combustor flame stability; (d) reducing lean blowout limit of
combustor; and (e) reducing combustion emissions.
[0035] Exemplary embodiments of methods, systems, and apparatus for
combustor flame stabilization are not limited to the specific
embodiments described herein, but rather, components of the systems
and/or steps of the methods may be utilized independently and
separately from other components and/or steps described herein. For
example, the methods may also be used in combination with other
systems requiring flame stabilization, and the associated methods,
and are not limited to practice with only the systems and methods
as described herein. Rather, the exemplary embodiment can be
implemented and utilized in connection with many other
applications, equipment, and systems that may benefit from thermal
control.
[0036] Although specific features of various embodiments of the
disclosure may be shown in some drawings and not in others, this is
for convenience only. In accordance with the principles of the
disclosure, any feature of a drawing may be referenced and/or
claimed in combination with any feature of any other drawing.
[0037] This written description uses examples to disclose the
embodiments, including the best mode, and also to enable any person
skilled in the art to practice the embodiments, including making
and using any devices or systems and performing any incorporated
methods. The patentable scope of the disclosure is defined by the
claims, and may include other examples that occur to those skilled
in the art. Such other examples are intended to be within the scope
of the claims if they have structural elements that do not differ
from the literal language of the claims, or if they include
equivalent structural elements with insubstantial differences from
the literal language of the claims.
* * * * *