U.S. patent application number 10/166960 was filed with the patent office on 2004-06-03 for gas turbine engine combustor can with trapped vortex cavity.
Invention is credited to Burrus, David Louis, Feitelberg, Alan S., Haynes, Joel Meier, Joshi, Narendra Digamber.
Application Number | 20040103663 10/166960 |
Document ID | / |
Family ID | 29583747 |
Filed Date | 2004-06-03 |
United States Patent
Application |
20040103663 |
Kind Code |
A1 |
Haynes, Joel Meier ; et
al. |
June 3, 2004 |
GAS TURBINE ENGINE COMBUSTOR CAN WITH TRAPPED VORTEX CAVITY
Abstract
A gas turbine engine combustor can downstream of a pre-mixer has
a pre-mixer flowpath therein and circumferentially spaced apart
swirling vanes disposed across the pre-mixer flowpath. A primary
fuel injector is positioned for injecting fuel into the pre-mixer
flowpath. A combustion chamber surrounded by an annular combustor
liner disposed in supply flow communication with the pre-mixer. An
annular trapped dual vortex cavity located at an upstream end of
the combustor liner is defined between an annular aft wall, an
annular forward wall, and a circular radially outer wall formed
therebetween. A cavity opening at a radially inner end of the
cavity is spaced apart from the radially outer wall. Air injection
first holes are disposed through the forward wall and air injection
second holes are disposed through the aft wall. Fuel injection
holes are disposed through at least one of the forward and aft
walls.
Inventors: |
Haynes, Joel Meier;
(Niskayuna, NY) ; Feitelberg, Alan S.; (Niskayuna,
NY) ; Burrus, David Louis; (Cincinnati, OH) ;
Joshi, Narendra Digamber; (Cincinnati, OH) |
Correspondence
Address: |
STEVEN J. ROSEN, PATENT ATTORNEY
4729 CORNELL RD.
CINCINNATI
OH
45241
US
|
Family ID: |
29583747 |
Appl. No.: |
10/166960 |
Filed: |
June 11, 2002 |
Current U.S.
Class: |
60/737 ;
60/750 |
Current CPC
Class: |
F23R 3/346 20130101;
F23R 2900/00015 20130101; F23R 3/283 20130101; F23R 3/286 20130101;
F23D 2900/14004 20130101; F23C 2900/07001 20130101; F23R 3/58
20130101 |
Class at
Publication: |
060/737 ;
060/750 |
International
Class: |
F23R 003/30 |
Claims
What is claimed is:
1. A gas turbine engine combustor can assembly comprising: a
combustor can downstream of a pre-mixer; said pre-mixer having a
pre-mixer upstream end, a pre-mixer downstream end and a pre-mixer
flowpath therebetween, a plurality of circumferentially spaced
apart swirling vanes disposed across said pre-mixer flowpath
between said upstream and downstream ends, and a primary fuel
injection means for injecting fuel into said pre-mixer flowpath;
said combustor can having a combustion chamber surrounded by an
annular combustor liner disposed in supply flow communication with
said pre-mixer; an annular trapped dual vortex cavity located at
said upstream end of said combustor liner and defined between an
annular aft wall, an annular forward wall, and a circular radially
outer wall formed therebetween; a cavity opening at a radially
inner end of said cavity spaced apart from said radially outer wall
and extending between said aft wall and said forward wall; air
injection first holes in said forward wall and air injection second
holes in said aft wall, said air injection first and second holes
spaced radially apart; and fuel injection holes in at least one of
said forward and aft walls.
2. A combustor can assembly as claimed in claim 1, further
comprising angled film cooling apertures disposed through said aft
wall, said forward wall, said and outer wall.
3. A combustor can assembly as claimed in claim 2, further
comprising said film cooling apertures through said aft walls are
angled radially outwardly, said film cooling apertures through said
forward walls are angled radially inwardly in a downstream
direction, and said film cooling apertures through said outer wall
are angled axially forwardly.
4. A combustor can assembly as claimed in claim 2, further
comprising said film cooling apertures through said aft walls are
angled radially inwardly, said film cooling apertures through said
forward walls are angled radially outwardly in a downstream
direction, and said film cooling apertures through said outer wall
are angled axially aftwardly.
5. A combustor can assembly as claimed in claim 2, wherein each of
said fuel injection holes is surrounded by a plurality of said air
injection second holes and said air injection first holes are
singularly arranged in a circumferential row.
6. A combustor can assembly as claimed in claim 5, further
comprising angled film cooling apertures disposed through said aft
wall, said forward wall, said and outer wall.
7. A combustor can assembly as claimed in claim 6, further
comprising said film cooling apertures through said aft walls are
angled radially outwardly, said film cooling apertures through said
forward walls are angled radially inwardly in a downstream
direction, and said film cooling apertures through said outer wall
are angled axially forwardly.
8. A combustor can assembly as claimed in claim 6, further
comprising said film cooling apertures through said aft walls are
angled radially inwardly, said film cooling apertures through said
forward walls are angled radially outwardly in a downstream
direction, and said film cooling apertures through said outer wall
are angled axially aftwardly.
9. A combustor can assembly as claimed in claim 1, wherein said
primary fuel injection means includes fuel cavities within said
swirling vanes, fuel injection holes extending through trailing
edges of said swirling vanes from the fuel cavities to said
pre-mixer flowpath.
10. A combustor can assembly as claimed in claim 9, further
comprising angled film cooling apertures disposed through said aft
wall, said forward wall, said and outer wall.
11. A combustor can assembly as claimed in claim 10, further
comprising said film cooling apertures through said aft walls are
angled radially outwardly, said film cooling apertures through said
forward walls are angled radially inwardly in a downstream
direction, and said film cooling apertures through said outer wall
are angled axially forwardly.
12. A combustor can assembly as claimed in claim 10, further
comprising said film cooling apertures through said aft walls are
angled radially inwardly, said film cooling apertures through said
forward walls are angled radially outwardly in a downstream
direction, and said film cooling apertures through said outer wall
are angled axially aftwardly.
13. A combustor can assembly as claimed in claim 10, wherein each
of said fuel injection holes is surrounded by a plurality of said
air injection second holes and said air injection first holes are
singularly arranged in a circumferential row.
14. A combustor can assembly as claimed in claim 13, further
comprising angled film cooling apertures disposed through said aft
wall, said forward wall, said and outer wall.
15. A combustor can assembly as claimed in claim 14, further
comprising said film cooling apertures through said aft walls are
angled radially outwardly, said film cooling apertures through said
forward walls are angled radially inwardly in a downstream
direction, and said film cooling apertures through said outer wall
are angled axially forwardly.
16. A combustor can assembly as claimed in claim 14, further
comprising said film cooling apertures through said aft walls are
angled radially inwardly, said film cooling apertures through said
forward walls are angled radially outwardly in a downstream
direction, and said film cooling apertures through said outer wall
are angled axially aftwardly.
17. A combustor can assembly as claimed in claim 1, further
comprising: a reverse flow combustor flowpath including, in
downstream serial flow relationship, an aft to forward portion
between an outer flow sleeve and said annular combustor liner, a
180 degree bend forward of said vortex cavity, and said pre-mixer
flowpath at a downstream end of said combustor flowpath; said
swirling vanes 32 disposed across said pre-mixer flowpath defined
between an outer flow sleeve and an inner flow sleeve.
18. A combustor can assembly as claimed in claim 17, further
comprising: said film cooling apertures through said aft walls are
angled radially inwardly, said film cooling apertures through said
forward walls are angled radially outwardly in a downstream
direction, said film cooling apertures through said outer wall are
angled axially aftwardly, said fuel injection holes and said air
injection second holes are disposed through said forward wall, and
said air injection first holes are disposed through said aft
wall.
19. A combustor can assembly as claimed in claim 18, wherein said
primary fuel injection means includes fuel cavities within said
swirling vanes, fuel injection holes extending through trailing
edges of said swirling vanes from the fuel cavities to said
pre-mixer flowpath.
20. A combustor can assembly as claimed in claim 18, further
comprising angled film cooling apertures disposed through said aft
wall, said forward wall, said and outer wall.
21. A combustor can assembly as claimed in claim 18, wherein each
of said fuel injection holes is surrounded by a plurality of said
air injection second holes and said air injection first holes are
singularly arranged in a circumferential row.
22. A combustor can assembly as claimed in claim 2, further
comprising a second stage pre-mixing convoluted mixer located
between said pre-mixer and said vortex cavity and including
circumferentially alternating lobes extending radially inwardly
into said pre-mixer flowpath.
23. A combustor can assembly as claimed in claim 22, further
comprising angled film cooling apertures disposed through said aft
wall, said forward wall, said and outer wall.
24. A combustor can assembly as claimed in claim 23, further
comprising: said film cooling apertures through said aft walls are
angled radially outwardly, said film cooling apertures through said
forward walls are angled radially inwardly in a downstream
direction, said film cooling apertures through said outer wall are
angled axially forwardly, said fuel injection holes and said air
injection second holes are disposed through said aft wall, and said
air injection first holes are disposed through said forward
wall.
25. A combustor can assembly as claimed in claim 24, wherein each
of said fuel injection holes is surrounded by a plurality of said
air injection second holes and said air injection first holes are
singularly arranged in a circumferential row.
Description
BACKGROUND OF THE INVENTION
[0001] This Invention was made with Government support under
Contract No. DE-FC26-01NT41020 awarded by the Department of Energy.
The Government has certain rights in this invention.
[0002] The present invention relates to gas turbine engine
combustors and, more particularly, to can-annular combustors with
pre-mixers.
[0003] Industrial gas turbine engines include a compressor for
compressing air that is mixed with fuel and ignited in a combustor
for generating combustion gases. The combustion gases flow to a
turbine that extracts energy for driving a shaft to power the
compressor and produces output power for powering an electrical
generator, for example. Electrical power generating gas turbine
engines are typically operated for extended periods of time and
exhaust emissions from the combustion gases are a concern and are
subject to mandated limits. Thus, the combustor is designed for low
exhaust emissions operation and, in particular, for low NOx
operation. A typical low NOx combustor includes a plurality of
combustor cans circumferentially adjoining each other around the
circumference of the engine. Each combustor can has a plurality of
pre-mixers joined to the upstream end. Lean burning pre-mixed low
NOx combustors have been designed to produce low exhaust emissions
but are susceptible to combustion instabilities in the combustion
chamber.
[0004] Diatomic nitrogen rapidly disassociates at temperatures
exceeding about 3000.degree. F. and combines with oxygen to produce
unacceptably high levels of NOx emissions. One method commonly used
to reduce peak temperatures and, thereby, reduce NOx emissions, is
to inject water or steam into the combustor. However, water/steam
injection is a relatively expensive technique and can cause the
undesirable side effect of quenching carbon monoxide (CO) burnout
reactions. Additionally, water/steam injection methods are limited
in their ability to reach the extremely low levels of pollutants
required in many localities. Lean pre-mixed combustion is a much
more attractive method of lowering peak flame temperatures and,
correspondingly, NOx emission levels. In lean pre-mixed combustion,
fuel and air are pre-mixed in a pre-mixing section and the fuel-air
mixture is injected into a combustion chamber where it is burned.
Due to the lean stoichiometry resulting from the pre-mixing, lower
flame temperatures and NOx emission levels are achieved. Several
types of low NOx emission combustors are currently employing lean
pre-mixed combustion for gas turbines, including can-annular and
annular type combustors.
[0005] Can-annular combustors typically consist of a cylindrical
can-type liner inserted into a transition piece with multiple
fuel-air pre-mixers positioned at the head end of the liner.
Annular combustors are also used in many gas turbine applications
and include multiple pre-mixers positioned in rings directly
upstream of the turbine nozzles in an annular fashion. An annular
burner has an annular cross-section combustion chamber bounded
radially by inner and outer liners while a can burner has a
circular cross-section combustion chamber bounded radially by a
single liner.
[0006] Industrial gas turbine engines typically include a combustor
designed for low exhaust emissions operation and, in particular,
for low NOx operation. Low NOx combustors are typically in the form
of a plurality of combustor cans circumferentially adjoining each
other around the circumference of the engine, with each combustor
can having a plurality of pre-mixers joined to the upstream ends
thereof. Each pre-mixer typically includes a cylindrical duct in
which is coaxially disposed a tubular centerbody extending from the
duct inlet to the duct outlet where it joins a larger dome defining
the upstream end of the combustor can and combustion chamber
therein.
[0007] A swirler having a plurality of circumferentially spaced
apart vanes is disposed at the duct inlet for swirling compressed
air received from the engine compressor. Disposed downstream of the
swirler are suitable fuel injectors typically in the form of a row
of circumferentially spaced-apart fuel spokes, each having a
plurality of radially spaced apart fuel injection orifices which
conventionally receive fuel, such as gaseous methane, through the
centerbody for discharge into the pre-mixer duct upstream of the
combustor dome.
[0008] The fuel injectors are disposed axially upstream from the
combustion chamber so that the fuel and air has sufficient time to
mix and pre-vaporize. In this way, the pre-mixed and pre-vaporized
fuel and air mixture support cleaner combustion thereof in the
combustion chamber for reducing exhaust emissions. The combustion
chamber is typically imperforate to maximize the amount of air
reaching the pre-mixer and, therefore, producing lower quantities
of NOx emissions and thus is able to meet mandated exhaust emission
limits.
[0009] Lean pre-mixed low NOx combustors are more susceptible to
combustion instability in the combustion chamber which causes the
fuel and air mixture to vary, thus, lowering the effectiveness of
the combustor to reduce emissions. Lean burning low NOx emission
combustors with pre-mixers are subject to combustion instability
that imposes serious limitations upon the operability of pre-mixed
combustion systems. There exists a need in the art to provide
combustion stability for a combustor which uses pre-mixing.
BRIEF SUMMARY OF THE INVENTION
[0010] A gas turbine engine combustor can assembly includes a
combustor can downstream of a pre-mixer having a pre-mixer upstream
end, a pre-mixer downstream end, and a pre-mixer flowpath
therebetween. A plurality of circumferentially spaced apart
swirling vanes are disposed across the pre-mixer flowpath between
the upstream and downstream ends. A primary fuel injector is used
for injecting fuel into the pre-mixer flowpath. The combustor can
has a combustion chamber surrounded by an annular combustor liner
disposed in supply flow communication with the pre-mixer. An
annular trapped dual vortex cavity is located at an upstream end of
the combustor liner and is defined between an annular aft wall, an
annular forward wall, and a circular radially outer wall formed
therebetween. A cavity opening at a radially inner end of the
cavity is spaced apart from the radially outer wall and extends
between the aft wall and the forward wall. Air injection first
holes are disposed through the forward wall and air injection
second holes are disposed through the aft wall. The air injection
first and second holes are spaced radially apart and fuel injection
holes are disposed through at least one of the forward and aft
walls.
[0011] An exemplary embodiment of the combustor can assembly
includes angled film cooling apertures disposed through the aft
wall angled radially outwardly in the downstream direction, film
cooling apertures disposed through the forward wall angled radially
inwardly, and film cooling apertures disposed through the outer
wall angled axially forwardly. Alternatively, the film cooling
apertures through the aft wall are angled radially inwardly in the
downstream direction, the film cooling apertures through the
forward wall are angled radially outwardly in the downstream
direction, and the film cooling apertures through the outer wall
are angled axially aftwardly. Each of the fuel injection holes is
surrounded by a plurality of the air injection second holes and the
air injection first holes are singularly arranged in a
circumferential row. The primary fuel injector includes fuel
cavities within the swirling vanes and fuel injection holes
extending through trailing edges of the swirling vanes from the
fuel cavities to the pre-mixer flowpath.
[0012] One alternative combustor can assembly has a reverse flow
combustor flowpath including, in downstream serial flow
relationship, an aft to forward portion between an outer flow
sleeve and the annular combustor liner, a 180 degree bend forward
of the vortex cavity, and the pre-mixer flowpath at a downstream
end of the combustor flowpath. The swirling vanes are disposed
across the pre-mixer flowpath defined between an outer flow sleeve
and an inner flow sleeve. Another alternative combustor can
assembly has a second stage pre-mixing convoluted mixer located
between the pre-mixer and the vortex cavity. The convoluted mixer
includes circumferentially alternating lobes extending radially
inwardly into the pre-mixer flowpath.
BRIEF DESCRIPTION OF THE DRAWINGS
[0013] While the specification concludes with claims particularly
pointing out and distinctly claiming the present invention, it is
believed that the same will be better understood from the following
description taken in conjunction with the accompanying drawings in
which:
[0014] FIG. 1 is a schematic illustration of a portion of an
industrial gas turbine engine having a low NOx pre-mixer and can
combustor with a trapped vortex cavity in accordance with an
exemplary embodiment of the present invention.
[0015] FIG. 2 is an enlarged longitudinal cross-sectional view
illustration of the can combustor illustrated in FIG. 1.
[0016] FIG. 3 is an enlarged longitudinal cross-sectional view
illustration of the trapped vortex cavity illustrated in FIG.
2.
[0017] FIG. 4 is an elevated view illustration taken in a direction
along 4-4 in FIG. 3.
[0018] FIG. 5 is a longitudinal cross-sectional view schematic
illustration of a first alternative can combustor with a convoluted
mixer between the pre-mixer and the can combustor.
[0019] FIG. 6 is an elevated view illustration of the convoluted
mixer taken in a direction along 6-6 in FIG. 5.
[0020] FIG. 7 is a longitudinal cross-sectional view schematic
illustration of a second alternative can combustor with a reverse
flow flowpath.
[0021] FIG. 8 is a longitudinal cross-sectional view illustration
of a fuel vane in the reverse flow flowpath through 8-8 in FIG.
7.
[0022] FIG. 9 is an enlarged view illustration of the trapped
vortex cavity illustrated in FIG. 8.
DETAILED DESCRIPTION OF THE INVENTION
[0023] Illustrated in FIG. 1 is an exemplary industrial gas turbine
engine 10 including a multi-stage axial compressor 12 disposed in
serial flow communication with a low NOx combustor 14 and a single
or multi-stage turbine 16. The turbine 16 is drivingly connected to
compressor 12 by a drive shaft 18 which is also used to drive an
electrical generator (not shown) for generating electrical power.
During operation, the compressor 12 discharges compressed air 20 in
a downstream direction D into the combustor 14 wherein the
compressed air 20 is mixed with fuel 22 and ignited for generating
combustion gases 24 from which energy is extracted by the turbine
16 for rotating the shaft 18 to power compressor 12 and driving the
generator or other suitable external load. The combustor 14 is
can-annular having a plurality of combustor can assemblies 25
circumferentially disposed about an engine centerline 4.
[0024] Referring further to FIG. 2, each of the combustor can
assemblies 25 includes a combustor can 23 directly downstream of a
pre-mixer 28 that forms a main air/fuel mixture in a fuel/air
mixture flow 35 in a pre-mixing zone 158 between the pre-mixer and
the combustor can. The combustor can 23 includes a combustion
chamber 26 surrounded by a tubular or annular combustor liner 27
circumscribed about a can axis 8 and attached to a combustor dome
29. The combustion chamber 26 has a body of revolution shape with
circular cross-sections normal to the can axis 8. In the exemplary
embodiment, the combustor liner 27 is imperforate to maximize the
amount of air reaching the pre-mixer 28 for reducing NOx emissions.
The generally flat combustor dome 29 is located at an upstream end
30 of the combustion chamber 26 and an outlet 31 is located at a
downstream end 33 of the combustion chamber. A transition section
(not illustrated) joins the plurality of combustor can outlets 31
to effect a common annular discharge to turbine 16.
[0025] The lean combustion process associated with the present
invention makes achieving and sustaining combustion difficult and
associated flow instabilities effect the combustors low NOx
emissions effectiveness. In order to overcome this problem within
combustion chamber 26, some technique for igniting the fuel/air
mixture and stabilizing the flame thereof is required. This is
accomplished by the incorporation of a trapped vortex cavity 40
formed in the combustor liner 27. The trapped vortex cavity 40 is
utilized to produce an annular rotating vortex 41 of a fuel and air
mixture as schematically depicted in the cavity in FIGS. 1, 2 and
3.
[0026] Referring to FIG. 3, an igniter 43 is used to ignite the
annular rotating vortex 41 of a fuel and air mixture and spread a
flame front into the rest of the combustion chamber 26. The trapped
vortex cavity 40 thus serves as a pilot to ignite the main air/fuel
mixture in the air/fuel mixture flow 35 that is injected into the
combustion chamber 26 from the air fuel pre-mixer 28. The trapped
vortex cavity 40 is illustrated as being substantially rectangular
in shape and is defined between an annular aft wall 44, an annular
forward wall 46, and a circular radially outer wall 48 formed
therebetween which is substantially perpendicular to the aft and
forward walls 44 and 46, respectively. The term "aft" refers to the
downstream direction D and the term "forward" refers to an upstream
direction U.
[0027] A cavity opening 42 extends between the aft wall 44 and the
forward wall 46 at a radially inner end 39 of the cavity 40, is
open to combustion chamber 26, and is spaced radially apart and
inwardly of the outer wall 48. In the exemplary embodiment
illustrated herein, the vortex cavity 40 is substantially
rectangular in cross-section and the aft wall 44, the forward wall
46, and the outer wall 48 are approximately equal in length in an
axially extending cross-section as illustrated in the FIGS.
[0028] Referring to FIG. 3 in particular, vortex driving aftwardly
injected air 110 is injected through air injection first holes 112
in the forward wall 46 positioned radially along the forward wall
positioned radially near the opening 42 at the radially inner end
39 of the cavity 40. Vortex driving forwardly injected air 116 is
injected through air injection second holes 114 in the aft wall 44
positioned radially near the outer wall 48. Vortex fuel 115 is
injected through fuel injection holes 70 in the aft wall 44 near
the radially outer wall 48. Each of the fuel injection holes 70 are
surrounded by several of the second holes 114 that are arranged in
a circular pattern. The first holes 112 in the forward wall 46 are
arranged in a singular circumferential row around the can axis 8 as
illustrated in FIG. 4. However, other arrangements may be used
including more than one row of the fuel injection holes 70 and/or
the first holes 112.
[0029] Referring to FIG. 3, the vortex fuel 115 enters trapped
vortex cavity 40 through a fuel injectors 68, which are centered
within the fuel injection holes 70. The fuel injector 68 is in flow
communication with an outer fuel manifold 74 that receives the
vortex fuel 115 by way of a fuel conduit 72. In the exemplary
embodiment of the invention, the fuel manifold 74 has an insulating
layer 80 in order to protect the fuel manifold from heat and the
insulating layer may contain either air or some other insulating
material.
[0030] Film cooling means, in the form of cooling apertures 84,
such as cooling holes or slots angled through walls, are well known
in the industry for cooling walls in the combustor. In the
exemplary embodiment of the invention, film cooling apertures 84
disposed through the aft wall 44, the forward wall 46, and the
outer wall 48 are used as the film cooling means. The film cooling
apertures 84 are angled to help promote the vortex 41 of fuel and
air formed within cavity 40 and are also used to cool the walls.
The film cooling apertures 84 are angled to flow cooling air 102 in
the direction of rotation 104 of the vortex. Due to the entrance of
air in cavity 40 from the first and second holes 112 and 114 and
the film cooling apertures 84, a tangential direction of the
trapped vortex 41 at the cavity opening 42 of the vortex cavity 40
is downstream D, the same as that of the fuel/air mixture entering
combustion chamber 26. This means that for a downstream D
tangential direction of the trapped vortex 41 at the cavity opening
42 of the vortex cavity 40, the film cooling apertures 84 through
the aft wall 44 are angled radially outwardly RO in the downstream
direction D, the film cooling apertures 84 through the forward wall
46 are angled radially inwardly RI, and the film cooling apertures
84 through the outer wall 48 are angled axially forwardly AF. For
an upstream U tangential direction of the trapped vortex 41 at the
cavity opening 42 of the vortex cavity 40 of the vortex 41, the
film cooling apertures 84 through the aft wall 44 are angled
radially inwardly RI in the downstream direction D, the film
cooling apertures 84 through the forward wall 46 are angled
radially outwardly RO in the downstream direction D, and the film
cooling apertures 84 through the outer wall 48 are angled axially
aftwardly AA (see FIGS. 7 and 9).
[0031] Accordingly, the combustion gases generated by the trapped
vortex within cavity 40 serves as a pilot for combustion of air and
fuel mixture received into the combustion chamber 26 from the
pre-mixer. The trapped vortex cavity 40 provides a continuous
ignition and flame stabilization source for the fuel/air mixture
entering combustion chamber 26. Since the trapped vortex performs
the flame stabilization function, it is not necessary to generate
hot gas recirculation zones in the main stream flow, as is done
with all other low NOx combustors. This allows a swirl-stabilized
recirculation zone to be eliminated from a main stream flow field
in the can combustor. The primary fuel would be injected into a
high velocity stream entering the combustion chamber without flow
separation or recirculation and with minimal risk of auto-ignition
or flashback and flame holding in the region of the fuel/air
pre-mixer.
[0032] A trapped vortex combustor can achieve substantially
complete combustion with substantially less residence time than a
conventional lean pre-mixed industrial gas turbine combustor. By
keeping the residence time in the combustion chamber relatively
short, the time spent at temperatures above the thermal NOx
formation threshold can be reduced, thus, reducing the amount of
Nox produced. A risk to this approach is increased CO levels due to
reduced time for complete CO burnout. However, it is believed that
the flame zone of the combustion chamber is very short due to
intense mixing between the vortex and the main air. The trapped
vortex provides high combustor efficiency under much shorter
residence time than conventional aircraft combustors. It is
expected that CO levels will be a key contributor to determination
of optimal combustor length and residence time.
[0033] Ignition, acceleration, and low-power operation would be
accomplished with fuel supplied only to the trapped vortex. At some
point in the load range, fuel would be introduced into the main
stream pre-mixer. Radially inwardly flow of hot combustion products
from the trapped vortex into the main stream would cause main
stream ignition. As load continued to increase, main stream fuel
injection would be increase and the trapped vortex fuel would be
decreased at a slower rate, such that combustor exit temperature
would rise. At full-load conditions, trapped vortex fuel flow would
be reduced to the point that the temperature in the vortex would be
below the thermal NOx formation threshold level, yet, still
sufficient to stabilize the main stream combustion. With the
trapped vortex running too lean to produce much thermal NOx and the
main stream residence time at high temperature too short to produce
much thermal NOx, the total emissions of the combustor would be
minimized.
[0034] In the exemplary embodiment illustrated herein the combustor
liner 27 includes a radially outerwardly opening annular cooling
slot 120 that is parallel to the aft wall 44 and operable to direct
and flow cooling air 102 along the aft wall 44. The combustor liner
27 includes a downstream opening annular cooling slot 128 is
operable to direct and flow cooling air 102 downstream along the
combustor liner 27 downstream of the cavity 40. The radially
outerwardly opening cooling slot 120 and the downstream opening
cooling slot 128 are parts of what is referred to as a cooling
nugget 117.
[0035] Referring again to FIG. 2, the pre-mixer 28 includes an
annular swirler 126 having a plurality of swirling vanes 32
circumferentially disposed about a hollow centerbody 45 across a
pre-mixer flowpath 134 which extends through a pre-mixer tube 140.
A fuel line 59 supplies fuel 22 to a fuel injector exemplified by
fuel cavities 130 within the swirling vanes 32 (see FIG. 8) of the
annular swirler 126. The fuel 22 is injected into the pre-mixer
flowpath 134 through fuel injection holes 132 which extend through
trailing edges 133 of the swirling vanes 32 from the fuel cavities
130 to the pre-mixer flowpath. An example of such a swirling vane
32 is illustrated in cross-section in FIG. 8. This is one primary
fuel injection means for injecting fuel into the pre-mixer flowpath
134. Other means are well known in the art and include, but are not
limited to, radially extending fuel rods that inject fuel in a
downstream direction in the pre-mixer flowpath 134 and central fuel
tubes that inject fuel radially into the pre-mixer flowpath 134.
The pre-mixer tube 140 is connected to the combustor dome 29 and
terminates at a pre-mixer nozzle 144 between the pre-mixer and the
combustion chamber 26. The hollow centerbody 45 is capped by an
effusion cooled centerbody tip 150.
[0036] Illustrated in FIG. 5 is a two stage pre-mixer 152 wherein a
first pre-mixing stage 157 includes the annular swirler 126. The
swirling vanes 32 are circumferentially disposed about the hollow
centerbody 45 across the pre-mixer flowpath 134 within the
pre-mixer tube 140. The fuel line 59 supplies fuel to fuel cavities
130 within the swirling vanes 32 of the annular swirler 126 as
further illustrated in FIG. 8. Downstream of the annular swirler
126 is a second pre-mixing stage 161 in the form of a convoluted
mixer 154 located between the first pre-mixing stage 157 and the
vortex cavity 40. The convoluted mixer 154 includes
circumferentially alternating lobes 159 extending radially inwardly
into the pre-mixer flowpath 134 and the fuel/air mixture flow
35.
[0037] A pre-mixing zone 158 extends between the annular swirler
126 and the convoluted mixer 154. The lobes 159 of the convoluted
mixer 154 direct a first portion 156 of the fuel/air mixture flow
35 from the pre-mixing zone 158 radially inwardly along the lobes
159 as illustrated in FIGS. 5 and 6. A second portion 166 of the
fuel/air mixture flow 35 from the pre-mixing zone 158 passes
between the lobes 159. The convoluted mixer 154 generates low
pressure zones 170 in wakes immediately downstream of the lobes
159. This encourages gases in the vortex cavity 40 to penetrate
deep into the fuel/air mixture flow 35 to provide good piloting
ignition of the air/fuel mixture in a combustion zone 172
downstream of the vortex cavity 40 in the combustion chamber 26.
The convoluted mixer 154 provides rapid mixing the combustion gases
from the vortex cavity 40. Some of the vortex fuel 115 from the
fuel injection holes 70 in the aft wall 44 near the radially outer
wall 48 will impinge on the forward wall 46. This fuel flows
radially inwardly up to and along an aft facing surface of the
convoluted mixer 154 and gets entrained in the air/fuel mixture
flow 35. This provides more mixing of the air/fuel mixture. The
convoluted mixer 154 anchors and stabilizes a flame front of the
air/fuel mixture in the combustion zone 172 and provides a high
degree of flame stability.
[0038] Illustrated in FIG. 7 is a dry low NOx single stage
combustor 176 with a reverse flow combustor flowpath 178. The
combustor flowpath 178 includes, in downstream serial flow
relationship, an aft to forward portion 180 between an outer flow
sleeve 182 and the annular combustor liner 27, a 180 degree bend
181 forward of the vortex cavity 40, and the pre-mixer flowpath 134
at a downstream end 135 of the combustor flowpath 178. The swirling
vanes 32 of the pre-mixer 28 are disposed across the pre-mixer
flowpath 134 defined between outer flow sleeve 182 and an inner
flow sleeve 184. The fuel line 59 supplies fuel 22 to the fuel
cavities 130 within the swirling vanes 32 of the annular swirler
126. The fuel is injected into the pre-mixer flowpath 134 through
the fuel injection holes 132 extending through trailing edges 133
of the swirling vanes 32 from the fuel cavities 130 as illustrated
in cross-section in FIG. 8.
[0039] Vortex driving aftwardly injected air 110 is injected
through air injection first holes 112 in the aft wall 44. The first
holes 112 are positioned lengthwise near the opening 42 at the
radially inner end 39 of the cavity 40. Vortex driving forwardly
injected air 116 is injected through air injection second holes 114
in the forward wall 46. The second holes 114 are positioned
radially along the forward wall as close as possible to the outer
wall 48. Vortex fuel 115 is injected through fuel injection holes
70 in the forward aft wall 46 near the radially outer wall 48. Each
of the fuel injection holes 70 are surrounded by several of the
second holes 114 that are arranged in a circular pattern. The first
holes 112 in the aft wall 44 are arranged in a singular
circumferential row around the can axis 8 as illustrated in FIG.
4.
[0040] Due to the entrance of air in cavity 40 from the first and
second holes 112 and 114 and the film cooling apertures 84, a
tangential direction of the trapped vortex 41 at the cavity opening
42 of the vortex cavity 40 is upstream which is opposite the
downstream direction of the fuel/air mixture entering combustion
chamber 26. This further promotes mixing of the hot combustion
gases of the vortex 41.
[0041] Accordingly, the combustion gases generated by the trapped
vortex within cavity 40 serves as a pilot for combustion of air and
fuel mixture received into the combustion chamber 26 from the
pre-mixer. The trapped vortex cavity 40 provides a continuous
ignition and flame stabilization source for the fuel/air mixture
entering combustion chamber 26. Since the trapped vortex performs
the flame stabilization function, it is not necessary to generate
hot gas recirculation zones in the main stream flow, as is done
with all other low NOx combustors. The film cooling apertures
within the cavities are angled to flow cooling air 102 in the
rotational direction that the vortex is rotating. Due to the
entrance of air in cavity 40 from the first and second holes 112
and 114 and the film cooling apertures 84, a tangential direction
of the trapped vortex 41 at the cavity opening 42 of the vortex
cavity 40 is downstream, the same as that of the fuel/air mixture
entering combustion chamber 26.
[0042] Since the primary fuel would be injected into a high
velocity stream through the swirler vanes with no flow separation
or recirculation, the risk of auto-ignition or flashback and flame
holding in the fuel/air pre-mixing region is minimized. It appears
that a trapped vortex combustor can is able to achieve complete
combustion with substantially less residence time than a
conventional lean pre-mixed industrial gas turbine combustor. By
keeping the residence time between the plane of the trapped vortex
and the exit of the combustor can relatively short, the time spent
at temperatures above the thermal NOx formation threshold can be
reduced.
[0043] While there have been described herein what are considered
to be preferred and exemplary embodiments of the present invention,
other modifications of the invention shall be apparent to those
skilled in the art from the teachings herein and, it is therefore,
desired to be secured in the appended claims all such modifications
as fall within the true spirit and scope of the invention.
[0044] Accordingly, what is desired to be secured by Letters Patent
of the United States is the invention as defined and differentiated
in the following claims:
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