U.S. patent application number 09/557178 was filed with the patent office on 2001-12-13 for premixing dry low nox emissions combustor with lean direct injection of gas fuel.
Invention is credited to Beebe, Kenneth W..
Application Number | 20010049932 09/557178 |
Document ID | / |
Family ID | 26942860 |
Filed Date | 2001-12-13 |
United States Patent
Application |
20010049932 |
Kind Code |
A1 |
Beebe, Kenneth W. |
December 13, 2001 |
Premixing dry low NOx emissions combustor with lean direct
injection of gas fuel
Abstract
Lean premixed combustion of a hydrocarbon fuel and air is
combined with lean direct injection of hydrocarbon fuel and carrier
fluid such as air or inert gas or a mixture of air and inert gas
into a combustor downstream of the premixed reaction zone in order
to achieve extremely low levels of emissions of oxides of nitrogen
at the high combustor exit temperatures required by advanced heavy
duty industrial gas turbines. One or more premixing fuel nozzles
are used to supply a lean mixture of hydrocarbon fuel and air to
the main or primary reaction zone of a gas turbine combustor. This
lean fuel/air mixture has an adiabatic flame temperature below the
temperature that would result in substantial thermal NOx formation.
After this low temperature reaction has been completed. Additional
fuel and carrier fluid are injected into the products of combustion
downstream of the main reaction zone in order to raise the
temperature of the mixture to the level required to operate an
advanced, high efficiency, heavy duty industrial gas turbine at
high load. Formation of nitrogen oxides in the region after this
secondary fuel and carrier fluid injection is minimized by partial
premixing of fuel and carrier fluid prior to ignition and by
minimizing the residence time between the secondary fuel injection
and the turbine first stage inlet.
Inventors: |
Beebe, Kenneth W.; (Galway,
NY) |
Correspondence
Address: |
Nixon & Vanderhye PC
1100 North Glebe Road 8th Floor
Arlington
VA
22201-4714
US
|
Family ID: |
26942860 |
Appl. No.: |
09/557178 |
Filed: |
April 21, 2000 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
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09557178 |
Apr 21, 2000 |
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09253009 |
Feb 19, 1999 |
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6192688 |
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09253009 |
Feb 19, 1999 |
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08643048 |
May 2, 1996 |
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6047550 |
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Current U.S.
Class: |
60/776 ;
60/737 |
Current CPC
Class: |
F23L 2900/07002
20130101; F23R 3/346 20130101; F23L 7/00 20130101; F23L 2900/07009
20130101; F23R 3/14 20130101; F23R 3/286 20130101; F23D 2900/14004
20130101 |
Class at
Publication: |
60/39.06 ;
60/737 |
International
Class: |
F23R 003/30 |
Claims
1. A combustor for a gas turbine comprising: a primary combustion
system for combusting a mixture of fuel and air in a primary
reaction zone, and operable in a plurality of gas turbine modes,
said gas turbine modes being determined based on a load range of
the gas turbine; and a secondary combustion system selectively
operable in a high load range mode of the plurality of gas turbine
modes, said secondary combustion system combusting a mixture of
fuel and carrier fluid in a secondary reaction zone.
2. A combustor according to claim 1, further comprising: a
combustor casing having an open end and an end cover assembly
secured to another end thereof; a flow sleeve mounted within said
casing; and a combustion liner within said flow sleeve and defining
at least said primary reaction zone; wherein said primary
combustion system comprises a sleeve cap assembly secured to said
casing and located axially downstream of said end cover assembly,
and at least one start-up fuel nozzle and a plurality of premixing
fuel nozzles communicating with said primary reaction zone.
3. A combustor according to claim 2, wherein each premixing fuel
nozzle comprises: a swirler including a plurality of swirl vanes
that impart rotation to entering air; and a plurality of fuel
spokes that distribute fuel in the rotating air stream.
4. A combustor according to claim 2, wherein said combustion liner
defines said secondary reaction zone downstream of said primary
reaction zone, said secondary combustion system comprising a lean
direct injection (LDI) fuel injector assembly communicating with
said secondary reaction zone.
5. A combustor according to claim 4, wherein said LDI fuel injector
assembly comprises an air manifold, a fuel manifold, and a
plurality of fuel/air injection spokes communicating with said air
manifold and said fuel manifold, said plurality of fuel/air
injection spokes penetrating the combustion liner for introducing
fuel and carrier fluid into said secondary reaction zone.
6. A combustor according to claim 5, wherein said carrier fluid is
air.
7. A combustor according to claim 5, wherein said carrier fluid is
one of inert gas or a mixture of air and inert gas.
8. A combustor according to claim 7, wherein said inert gas is one
of steam or nitrogen.
9. A combustor according to claim 1, wherein said secondary
combustion system comprises a lean direct injection (LDI) fuel
injector assembly.
10. A combustor according to claim 9, wherein said LDI fuel
injector assembly comprises an air manifold, a fuel manifold, and a
plurality of fuel/air injection spokes communicating with said air
manifold and said fuel manifold.
11. A combustor according to claim 1, further comprising a
transition piece disposed downstream of said primary combustion
system and said secondary combustion system for flowing hot gases
of combustion to turbine nozzles of the gas turbine.
12. A combustor according to claim 1, wherein said carrier fluid is
air.
13. A combustor according to claim 1, wherein said carrier fluid is
one of inert gas or a mixture of air and inert gas.
14. A combustor according to claim 13, wherein said inert gas is
one of steam or nitrogen.
15. A gas turbine comprising: a compressor section for pressurizing
inlet air; a combustion section disposed downstream of the
compressor section for receiving the pressurized inlet air; and a
turbine section disposed downstream of the combustion section for
receiving hot products of combustion from the combustion section,
wherein the combustion section comprises: a primary combustion
system for combusting a mixture of fuel and air in a primary
reaction zone, and operable in a plurality of gas turbine modes,
said gas turbine modes being determined based on a load range of
the gas turbine, and a secondary combustion system selectively
operable in a high load range mode of the plurality of gas turbine
modes, said secondary combustion system combusting a mixture of
fuel and carrier fluid in a secondary reaction zone.
16. A gas turbine according to claim 15, wherein said combustion
section further comprises: a combustor casing having an open end
and an end cover assembly secured to another end thereof; a flow
sleeve mounted within said casing; and a combustion liner within
said flow sleeve and defining at least a primary reaction zone;
wherein said primary combustion system comprises a sleeve cap
assembly secured to said casing and located axially downstream of
said end cover assembly, and at least one start-up fuel nozzle and
a plurality of premixing fuel nozzles communicating with said
primary reaction zone.
17. A gas turbine according to claim 16, wherein each premixing
fuel nozzle comprises: a swirler including a plurality of swirl
vanes that impart rotation to entering air; and a plurality of fuel
spokes that distribute fuel in the rotating air stream.
18. A gas turbine according to claim 16, wherein said combustion
liner defines said secondary reaction zone downstream of said
primary reaction zone, said secondary combustion system comprising
a lean direct injection (LDI) fuel injector assembly communicating
with said secondary reaction zone.
19. A gas turbine according to claim 18, wherein said LDI fuel
injector assembly comprises an air manifold, a fuel manifold, and a
plurality of fuel/air injection spokes communicating with said air
manifold and said fuel manifold, said plurality of fuel/air
injection spokes penetrating the combustion liner for introducing
fuel and carrier fluid into said secondary reaction zone.
20. A gas turbine according to claim 19, wherein said carrier fluid
is air.
21. A gas turbine according to claim 19, wherein said carrier fluid
is one of inert gas or a mixture of air and inert gas.
22. A gas turbine according to claim 21, wherein said inert gas is
one of steam or nitrogen.
23. A gas turbine according to claim 15, wherein said secondary
combustion system comprises a lean direct injection (LDI) fuel
injector assembly.
24. A gas turbine according to claim 23, wherein said LDI fuel
injector assembly comprises an air manifold, a fuel manifold, and a
plurality of fuel/air injection spokes communicating with said air
manifold and said fuel manifold.
25. A gas turbine according to claim 15, wherein said combustion
system further comprises a transition piece disposed downstream of
said primary combustion system and said secondary combustion system
for flowing hot gases of combustion to the turbine section.
26. A gas turbine according to claim 15, wherein said carrier fluid
is air.
27. A gas turbine according to claim 15, wherein said carrier fluid
is one of inert gas or a mixture of air and inert gas.
28. A gas turbine according to claim 27, wherein said inert gas is
one of steam or nitrogen.
29. A method of combustion in a gas turbine combustor including a
primary combustion system operable in a plurality of gas turbine
modes and including at least one start-up fuel nozzle and a
plurality of premixing fuel nozzles communicating with a primary
reaction zone, and a lean direct injection (LDI) combustion system
communicating with a secondary reaction zone, the method
comprising: (a) in a low-range turbine load mode, supplying fuel to
the said at least one start-up fuel nozzle and mixing the fuel with
air in the primary reaction zone; (b) in a mid-range turbine load
mode, supplying fuel to the premixing fuel nozzles and premixing
the fuel with air for combustion in the primary reaction zone; and
(c) in a high-range turbine load mode, carrying out step (b) and
then supplying fuel and carrier fluid to the LDI combustion system
for combustion in the secondary reaction zone.
30. The method of claim 29, wherein said carrier fluid is air.
31. The method of claim 29, wherein said carrier fluid is one of
inert gas and a mixture of air and inert gas.
32. The method of claim 31, wherein said inert gas is one of steam
or nitrogen.
33. A method of combustion for achieving low levels of emissions of
oxides of nitrogen (NOx) at high combustor exit temperatures
comprising lean direct injection of hydrocarbon fuel and carrier
fluid into a combustor downstream of a premixed combustion zone of
said combustor to which a lean mixture of hydrocarbon fuel and air
have been supplied and combusted at an adiabatic flame temperature
below the temperature that would result in substantial thermal NOx
formation.
34. The method of claim 33 in which the hydrocarbon fuel and
carrier fluid are partially premixed prior to lean direct injection
into the products of combustion from the premixed combustion zone,
thereby causing auto-ignition and raising the temperature of the
resulting combustion products.
35. The method of claim 33, wherein said carrier fluid is air.
36. The method of claim 33, wherein said carrier fluid is one of
inert gas or a mixture of air and inert gas.
37. The method of claim 36, wherein said inert gas is one of steam
or nitrogen.
Description
[0001] This application is a continuation-in-part of U.S. patent
application Ser. No. 08/643,048, filed May 2, 1996.
BACKGROUND OF THE INVENTION
[0002] This invention relates to gas and liquid fuel turbines and,
more specifically, to combustors in industrial gas turbines used in
power generation plants.
[0003] Gas turbine manufacturers, including General Electric, are
currently involved in research and engineering programs to produce
new gas turbines that will operate at high efficiency without
producing undesirable air polluting emissions. The primary air
polluting emissions usually produced by gas turbines burning
conventional hydrocarbon fuels are oxides of nitrogen, carbon
monoxide and unburned hydrocarbons. It is well known in the art
that oxidation of molecular nitrogen in air breathing engines is
highly dependent upon the maximum hot gas temperature in the
combustion system reaction zone and the residence time for the
reactants at the highest temperatures reached within the combustor.
The level of thermal NOx formation is minimized by maintaining the
reaction zone temperature below the level at which thermal NOx is
formed or by maintaining an extremely short residence time at high
temperature such that there is insufficient time for the NOx
formation reactions to progress.
[0004] One preferred method of controlling the temperature of the
reaction zone of a heat engine combustor below the level at which
thermal NOx is formed is to premix fuel and air to a lean mixture
prior to combustion. U.S. Pat. No. 4,292,801 dated October 1981,
the disclosure of which is hereby incorporated by reference,
describes a dual stage-dual mode low NOx combustor for gas turbine
application which is one of the pioneering combustor designs based
on lean premixed combustion technology. U.S. Pat. No. 5,259,184
dated November 1993, the disclosure of which is also hereby
incorporated by reference, describes a dry low NOx single stage
dual mode combustor construction for a gas turbine. The thermal
mass of the excess air present in the reaction zone of a lean
premixed combustor absorbs heat and reduces the temperature rise of
the products of combustion to a level where thermal NOx is not
formed. Even with this technology, for the most advanced high
efficiency heavy duty industrial gas turbines, the required
temperature of the products of combustion at the combustor
exit/first stage turbine inlet at maximum load is so high that the
combustor must be operated with peak gas temperature in the
reaction zone which exceeds the thermal NOx formation threshold
temperature resulting in significant NOx formation even though the
fuel and air are premixed lean. The problem to be solved is to
obtain combustor exit temperatures high enough to operate the most
advanced, high efficiency heavy duty industrial gas turbines at
maximum load without forming a significant amount of thermal
NOx.
[0005] Lean premixed combustion of hydrocarbon fuels in air is
widely used throughout the gas turbine industry as a method of
reducing air pollutant levels, in particular thermal NOx emissions
levels, for gas turbine combustors. Lean direct injection (LDI) of
hydrocarbon fuel and air has also been shown to be an effective
method for reducing NOx emission levels for gas turbine combustion
systems although not as effective as lean premixed combustion. An
example of an LDI fuel injector assembly is described in an article
from the 1987 Tokyo International Gas Turbine Congress entitled
"Lean Primary Zones: Pressure Loss and Residence Time Influences on
Combustion Performance and NOx Emissions," the disclosure of which
is hereby incorporated by reference. The present invention combines
these two technologies; i.e., lean premixed combustion and lean
direct fuel injection, in a novel and unique manner in order to
achieve extremely low air pollutant emissions levels, particularly
oxides of nitrogen, when operating an advanced, high efficiency,
heavy duty industrial gas turbine at high load.
BRIEF SUMMARY OF THE INVENTION
[0006] There is thus a particular need to combine premixed
combustion of a lean mixture of hydrocarbon fuel and air with lean
direct injection of hydrocarbon fuel and a carrier fluid such as
air or inert gas or a mixture of air and inert gas into the
products of lean premixed combustion late in the combustion
process, and thereby produce a combustion system that will yield
very low emissions of air pollutants, in particular oxides of
nitrogen, when operating an advanced, high efficiency, heavy duty
industrial gas turbine at high load. Moreover, this invention is
intended to accomplish this objective while operating the premixed
combustion reaction zone with a fuel/air mixture that is lean
enough to ensure that the thermal NOx formation in the reaction
zone is negligible and while operating the entire combustion system
at an overall fuel/air mixture strength that exceeds that of the
premixed reaction zone by the amount necessary to meet the inlet
temperature demands of the gas turbine. This invention is
particularly advantageous in applications where the inlet
temperature demands of the turbines are so high as to preclude the
possibility of achieving very low thermal NOx emissions levels by
lean premixed combustion alone.
[0007] These and other advantages are achieved by providing a
combustor for a gas turbine including a primary combustion system
operable in a plurality of gas turbine modes, the gas turbine modes
being determined based on a load range on the gas turbine, and a
secondary combustion system selectively operable in a high load
range mode of the plurality of gas turbine modes.
[0008] The combustor may further be provided with a combustor
casing having an open end and an end cover assembly secured to
another end thereof, a flow sleeve mounted within the casing, and a
combustion liner within the flow sleeve and defining at least a
primary reaction zone. The primary combustion system preferably
includes a sleeve cap assembly secured to the casing and located
axially downstream of the end cover assembly, and at least one
start up fuel nozzle and premixing fuel nozzles communicating with
the primary reaction zone. In this regard, each premixing fuel
nozzle preferably includes a swirler including a plurality of swirl
vanes that impart rotation to entering air, and a plurality of fuel
spokes that distribute fuel in the rotating air stream. The
combustion liner may also define a secondary reaction zone
downstream of the primary reaction zone. In this context, the
secondary combustion system includes a lean direct injection (LDI)
fuel injector assembly communicating with the secondary reaction
zone. The LDI fuel injector assembly preferably includes an air
manifold, a fuel manifold, and a plurality of fuel/air injection
spokes communicating with the air manifold and the fuel manifold.
The plurality of fuel/air injection spokes penetrate the combustion
liner and introduce fuel and carrier fluid into the secondary
reaction zone.
[0009] In accordance with another aspect of the invention, there is
provided a gas turbine including a compressor section that
pressurizes inlet air, a combustion section disposed downstream of
the compressor section that receives the pressurized inlet air, and
a turbine section disposed downstream of the combustion section and
receiving hot products of combustion from the combustion section.
The combustion section includes a circular array of
circumferentially spaced combustors according to the invention.
[0010] In accordance with still another aspect of the invention,
there is provided a method of combustion in a gas turbine combustor
according to the invention. The method includes the steps of (a) in
a low range turbine load mode, supplying fuel to start up fuel
nozzles and mixing the fuel with air in a primary reaction zone,
(b) in a mid-range turbine load mode, supplying fuel to premixing
fuel nozzles and premixing the fuel with air prior to entering the
primary reaction zone, and (c) in a high-range turbine load mode,
carrying out step (b) and then supplying secondary fuel and carrier
fluid to a secondary combustion system and introducing fuel and
carrier fluid into a secondary reaction zone.
BRIEF DESCRIPTION OF THE DRAWINGS
[0011] These and other aspects of the present invention will become
clear in the following description of the invention with reference
to the accompanying drawings in which:
[0012] FIG. 1 is a schematic cross-sectional illustration of a lean
premixed combustor forming part of a gas turbine and constructed in
accordance with the present invention;
[0013] FIG. 2 is a cross-sectional view thereof taken generally
along line 2-2 in FIG. 1; and
[0014] FIG. 3 is a cross-sectional illustration of one fuel/air
injection spoke taken from FIG. 2.
DETAILED DESCRIPTION OF THE INVENTION
[0015] Reference will now be made in detail to the present
preferred embodiments of the invention, an example of which is
illustrated in the accompanying drawings.
[0016] As is well known, a gas turbine includes a compressor
section, a combustion section and a turbine section. The compressor
section is driven by the turbine section through a common shaft
connection. The combustion section typically includes a circular
array of a plurality of circumferentially spaced combustors. A
fuel/air mixture is burned in each combustor to produce the hot
energetic flow of gas, which flows through a transition piece for
flowing the gas to the turbine blades of the turbine section. A
conventional combustor is described in the above-noted U.S. Pat.
No. 5,259,184. For purposes of the present description, only one
combustor is illustrated, it being appreciated that all of the
other combustors arranged about the turbine are substantially
identical to the illustrated combustor.
[0017] Referring now to FIG. 1, there is shown generally at 10, a
combustor for a gas turbine engine including a lean premixed
combustion assembly 12, a secondary or lean direct injection (LDI)
fuel injector assembly 50, and a transition piece 18 for flowing
hot gases of combustion to the turbine nozzles 11 and the turbine
blades (not shown). The lean premixed combustor assembly 12
includes a casing 20, an end cover 22, a plurality of start-up fuel
nozzles 24, a plurality of premixing fuel nozzles 14, a cap
assembly 30, a flow sleeve 17, and a combustion liner 28 within the
sleeve 17. A suitable cap assembly is described in U.S. Pat. No.
5,274,991, the disclosure of which is hereby incorporated by
reference. An ignition device (not shown) is provided and
preferably comprises an electrically energized spark plug.
Combustion in the lean premixed combustor assembly 12 occurs within
the combustion liner 28. Combustion air is directed within the
liner 28 via the flow sleeve 17 and enters the combustion liner
through a plurality of openings formed in the cap assembly 30. The
air enters the liner under a pressure differential across the cap
assembly 30 and mixes with fuel from the start-up fuel nozzles 24
and/or the premixing fuel nozzles 14 within the liner 28.
Consequently, a combustion reaction occurs within the liner 28
releasing heat for the purpose of driving the gas turbine. High
pressure air for the lean premixed combustor assembly 12 enters the
flow sleeve 17 and a transition piece impingement sleeve 15, from
an annular plenum 2. This high pressure air is supplied by a
compressor, which is represented by a series of vanes and blades at
13 and a diffuser 42.
[0018] Each premixing fuel nozzle 14 includes a swirler 4,
consisting of a plurality of swirl vanes that impart rotation to
the entering air and a plurality of fuel spokes 6 that distribute
fuel in the rotating air stream. The fuel and air then mix in an
annular passage within the premix fuel nozzle 14 before reacting
within the primary reaction zone 8.
[0019] The LDI fuel injector assembly 50 is provided for operating
at gas turbine high load conditions. Referring to FIGS. 2 and 3,
the assembly 50 includes an air manifold 51, a fuel manifold 52,
and a plurality of fuel/air injection spokes 53 that penetrate the
combustion liner 28 and introduce additional fuel and carrier fluid
into the secondary reaction zone 19 within the combustor assembly.
This secondary fuel/carrier fluid mixture is ignited by the hot
products of combustion exiting the primary reaction zone 8, and the
resulting secondary hydrocarbon fuel oxidation reactions go to
completion in the transition piece 18. The secondary fuel is
injected into the secondary carrier fluid via a plurality of fuel
orifices 57, and the combination of secondary fuel and secondary
carrier fluid is injected into the secondary reaction zone 19 via a
plurality of air orifices 56 in each fuel/air injection spoke
53.
[0020] In operation of the gas turbine, there are three distinct
operating modes depending upon the load range on the gas turbine.
The first operating mode is at low turbine load (about 0-30% of
base load) and during initial start up. In this mode, hydrocarbon
fuel is supplied to the start-up fuel nozzles 24, and combustion
air is provided to the liner 28 through the plurality of openings
in the cap assembly 30 for mixing with the fuel from the start-up
fuel nozzles 24. A diffusion flame reaction occurs within the
combustion liner 28 at the primary reaction zone 8. This reaction
is initiated by an electrically energized spark plug.
[0021] At mid-range operating conditions (about 30-80% of base
load), hydrocarbon fuel is supplied to the premixing fuel nozzles
14 via the fuel spokes 6. The premixer 14 mixes the hydrocarbon
fuel with air from the swirler 4, and the mixture enters the
primary reaction zone 8. The mixture of fuel and air ignites in the
presence of the diffusion flame from the start-up fuel nozzles 14.
Once the premixed combustion reaction has been initiated,
hydrocarbon fuel is diverted from the start-up fuel nozzles 24 to
the premixing fuel nozzles 14. The diffusion flame in the primary
reaction zone 8 then goes to extinction, and the combustion
reaction in the primary reaction zone 8 becomes entirely premixed.
Because the fuel/air mixture entering the primary reaction zone 8
is lean, the combustion reaction temperature is too low to produce
a significant amount of thermal NOx. The hydrocarbon fuel oxidation
reactions go to completion in the primary reaction zone 8 within
the combustion liner 28. Thus, during mid-range load conditions,
the temperature of the combustion reaction is too low to produce a
significant amount of thermal NOx.
[0022] Under high load conditions (about 80% of base load to peak
load), premixed combustion is carried out as described above.
Additionally, hydrocarbon fuel and carrier fluid are supplied to
the LDI fuel injector assembly 50. In preferred forms, the carrier
fluid can be air or an inert gas such as nitrogen or steam or a
mixture of air and inert gas. The assembly 50 introduces secondary
fuel and carrier fluid into the secondary reaction zone 19 where
auto-ignition occurs due to the high temperatures existing within
the combustion liner 28 at mid-load and high load conditions. The
secondary hydrocarbon fuel oxidation reactions go to completion in
the transition piece 18. Because the secondary fuel/carrier fluid
mixture entering the transition piece 18 is lean, the combustion
reaction temperature is lower than the stoichiometric flame
temperature, and the thermal NOx formation rate is low. Since the
residence time in the transition piece 18 is short and the thermal
NOx formation rate is low, very little thermal NOx is formed during
secondary fuel combustion.
[0023] Consequently, it will be appreciated that NOx emissions are
substantially minimized or eliminated through the mid-load and high
load operating ranges of high firing temperature, high efficiency
heavy duty industrial gas turbines. This has been accomplished
simply and efficiently and by a unique cooperation of essentially
known gas turbine elements. Both lean premixed combustion, used as
the primary combustion system for this invention, and lean direct
fuel injection, used as the secondary combustion system for this
invention, are well known NOx abatement methods in the gas turbine
industry. This invention is a novel and unique combination of these
methods to achieve extremely low NOx emission levels for state of
the art, high efficiency, heavy duty industrial gas turbines.
[0024] While the invention has been described in connection with
what is presently considered to be the most practical and preferred
embodiments, it is to be understood that the invention is not to be
limited to the disclosed embodiments, but on the contrary, is
intended to cover various modifications and equivalent arrangements
included within the spirit and scope of the appended claims.
* * * * *