U.S. patent application number 13/751301 was filed with the patent office on 2013-05-30 for system and method for reducing combustion dynamics in a turbomachine.
This patent application is currently assigned to GENERAL ELECTRIC COMPANY. The applicant listed for this patent is Fei Han, Vasanth Srinivasa Kothnur, Kapil Kumar Singh. Invention is credited to Fei Han, Vasanth Srinivasa Kothnur, Kapil Kumar Singh.
Application Number | 20130133331 13/751301 |
Document ID | / |
Family ID | 42111144 |
Filed Date | 2013-05-30 |
United States Patent
Application |
20130133331 |
Kind Code |
A1 |
Singh; Kapil Kumar ; et
al. |
May 30, 2013 |
SYSTEM AND METHOD FOR REDUCING COMBUSTION DYNAMICS IN A
TURBOMACHINE
Abstract
A turbomachine includes a combustion chamber, and at least one
pre-mixer mounted to the combustion chamber. The at least one
pre-mixer includes a main body having a first end portion that
extends to a second end portion. The first end portion is
configured to receive an amount of fuel and an amount of air and
the second end portion defines an exit plane from which a fuel-air
mixture discharges into the combustion chamber. The turbomachine
also includes a combustion dynamics reduction system operatively
coupled to the at least one pre-mixer. The combustion dynamics
reduction system includes at least one of a boundary layer
perturbation mechanism and an acoustic wave introduction system
which disrupt a flow pattern of the fuel-air mixture within the at
least one pre-mixer.
Inventors: |
Singh; Kapil Kumar;
(Rexford, NY) ; Kothnur; Vasanth Srinivasa;
(Clifton Park, NY) ; Han; Fei; (Clifton Park,
NY) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Singh; Kapil Kumar
Kothnur; Vasanth Srinivasa
Han; Fei |
Rexford
Clifton Park
Clifton Park |
NY
NY
NY |
US
US
US |
|
|
Assignee: |
GENERAL ELECTRIC COMPANY
Schenectady
NY
|
Family ID: |
42111144 |
Appl. No.: |
13/751301 |
Filed: |
January 28, 2013 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
12363955 |
Feb 2, 2009 |
|
|
|
13751301 |
|
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Current U.S.
Class: |
60/772 ;
60/725 |
Current CPC
Class: |
F23R 3/00 20130101; F23R
3/286 20130101; F23R 2900/00013 20130101 |
Class at
Publication: |
60/772 ;
60/725 |
International
Class: |
F23R 3/28 20060101
F23R003/28 |
Claims
1. A turbomachine comprising: a combustion chamber; at least one
pre-mixer mounted to the combustion chamber, the at least one
pre-mixer including a main body including a first end portion that
extends to a second end portion, the first end portion being
configured to receive an amount of fuel and an amount of air and
the second end portion defining an exit plane from which a fuel-air
mixture discharges into the combustion chamber; and a combustion
dynamics reduction system operatively coupled to the at least one
pre-mixer, the combustion dynamics reduction system including at
least one of a boundary layer perturbation mechanism and an
acoustic wave introduction system that disrupt a flow pattern of
the fuel-air mixture within the at least one pre-mixer.
2. The turbomachine according to claim 1, wherein the combustion
dynamics reduction system includes a boundary layer perturbation
mechanism, the boundary layer perturbation mechanism including one
of an air/inert injection system operatively coupled to the
pre-mixer and mechanical member mounted in the at least one
pre-mixer.
3. The turbomachine according to claim 2, wherein the boundary
layer perturbation mechanism includes an air/inert injection
system, the air/inert injection system including an inlet for
receiving an amount of air/inert and an outlet for releasing the
amount of air/inert.
4. The turbomachine according to claim 3, wherein the outlet is
positioned in the pre-mixer.
5. The turbomachine according to claim 4, wherein the outlet is
positioned adjacent the exit plane.
6. The turbomachine according to claim 2, wherein the boundary
layer perturbation system includes a mechanical member mounted in
the pre-mixer.
7. The turbomachine according to claim 6, wherein the mechanical
member comprises at least one protrusion mounted in the pre-mixer,
the at least one protrusion modifying the flow pattern of the
fuel-air mixture.
8. The turbomachine according to claim 7, wherein the at least one
protrusion is mounted adjacent the exit plane.
9. The turbomachine according to claim 1, wherein the combustion
dynamics reduction system comprises an acoustic wave introduction
system, the acoustic wave introduction system including an acoustic
driver operatively connected to the at least one pre-mixer.
10. The turbomachine according to claim 9, wherein the combustion
dynamics reduction system includes a fluid introduction system, the
fluid introduction system being operatively connected to the at
least one pre-mixer.
11. A method of reducing combustion dynamics in a turbomachine
comprising: directing a fuel-air mixture through a pre-mixer into a
combustion chamber, the pre-mixer; and reducing combustion dynamics
by disrupting a flow pattern of the fuel-air mixture within the at
least one pre-mixer.
12. The method of claim 11, wherein reducing combustion dynamics
comprises perturbing a boundary layer of the fuel-air mixture
passing from the pre-mixer.
13. The method of claim 12, wherein perturbing a boundary layer of
the fuel-air mixture comprises injecting air into the fuel-air
mixture at an exit plane of the pre-mixer.
14. The method of claim 12, wherein perturbing a boundary layer of
the fuel-air mixture comprises passing the fuel-air mixture over at
least one mechanical member arranged in the pre-mixer.
15. The method of claim 14, wherein passing the fuel-air mixture
over at least one mechanical member arranged in the pre-mixer
comprises passing the fuel-air mixture over at least one protrusion
mounted in the pre-mixer, the at least one protrusion modifying a
flow pattern of the fuel-air mixture.
16. The method of claim 15, wherein passing the fuel-air mixture
over at least one protrusion comprises passing the fuel-air mixture
over the at least one protrusion arranged at an exit plane of the
pre-mixer.
17. The method of claim 11, wherein reducing combustion dynamics
comprises perturbing a base portion of a flame at an exit plane of
the pre-mixer.
18. The method of claim 17, wherein perturbing the base portion of
the flame comprises introducing an acoustic wave into the
pre-mixer.
19. The method of claim 18, wherein introducing the acoustic wave
into the pre-mixer comprises varying a frequency of the acoustic
wave.
20. The method of claim 18, further comprising: introducing a fluid
into the pre-mixing to further perturb the base portion of the
flame.
Description
BACKGROUND OF THE INVENTION
[0001] The subject matter disclosed herein relates to the art of
turbomachines and, more particularly, to a system and method for
reducing combustion dynamics in a turbomachine.
[0002] Combustion dynamics are a phenomenon in gas turbomachines
utilizing lean pre-mixed combustion. Combustion dynamics include
low-frequency, longitudinal dynamics and high-frequency screech
caused by the excitation of radial and azimuthal modes of the
combustion chambers by the swirling flames. Both the low and high
frequencies include a combustion field component and an acoustic
component, that pass along the combustor during combustion. Under
certain operating conditions, the combustion component and the
acoustic component couple to create both low and high frequency
dynamic fields. The low and high frequency dynamic fields have a
negative impact on various turbomachine components. More
specifically, dynamic fields passing from the combustor may lead to
high cycle fatigue (HCF) for downstream turbomachine
components.
[0003] To address this problem, turbomachines are operated at less
than optimum levels, i.e., certain operating conditions are avoided
in order to avoid circumstances that are conducive to combustion
screech. While effective at reducing combustion dynamics, avoiding
these operating levels restricts an overall operating envelope of
the turbomachine.
[0004] Another approach to the problem of combustion dynamics is to
modify combustor input conditions. More specifically, fluctuations
in fuel-air ratio are known to cause combustion dynamics that lead
to combustion screech. Creating perturbations in the fuel-air
mixture by changing fuel flow rate can disengage the combustion
field from the acoustic field to suppress combustion screech. While
both of the above approaches are effective at reducing combustion
dynamics, avoiding various operating levels restricts an overall
operating envelope of the turbomachine while manipulating the
fuel-air ratio requires a coupled control scheme and may also lead
to less than efficient combustion.
BRIEF DESCRIPTION OF THE INVENTION
[0005] According to one aspect of the invention, a turbomachine
includes a combustion chamber, and at least one pre-mixer mounted
to the combustion chamber. The at least one pre-mixer includes a
main body having a first end portion that extends to a second end
portion. The first end portion is configured to receive an amount
of fuel and an amount of air and the second end portion defines an
exit plane from which a fuel-air mixture discharges into the
combustion chamber. The turbomachine also includes a combustion
dynamics reduction system operatively coupled to the at least one
pre-mixer. The combustion dynamics reduction system includes at
least one of a boundary layer perturbation mechanism and an
acoustic wave introduction system which disrupt a flow pattern of
the fuel-air mixture within the at least one pre-mixer.
[0006] According to another aspect of the invention, a method of
reducing combustion dynamics in a turbomachine includes directing a
fuel-air mixture through a pre-mixer into a combustion chamber, and
reducing combustion dynamics by disrupting a flow pattern of the
fuel-air mixture within the at least one pre-mixer.
[0007] These and other advantages and features will become more
apparent from the following description taken in conjunction with
the drawings.
BRIEF DESCRIPTION OF THE DRAWING
[0008] The subject matter which is regarded as the invention is
particularly pointed out and distinctly claimed in the claims at
the conclusion of the specification. The foregoing and other
features, and advantages of the invention are apparent from the
following detailed description taken in conjunction with the
accompanying drawings in which:
[0009] FIG. 1 is a partial, cross-sectional side view of a
turbomachine including a combustion dynamics reduction system in
accordance with exemplary embodiments of the invention;
[0010] FIG. 2 is a cross-sectional side view of a combustor portion
of the turbomachine of FIG. 1;
[0011] FIG. 3 is a schematic view of an injection nozzle assembly
including a combustion dynamics reduction system in accordance with
one exemplary aspect of the invention;
[0012] FIG. 4 is a schematic view of an injection nozzle assembly
including a combustion dynamics reduction system in accordance with
another exemplary aspect of the invention; and
[0013] FIG. 5 is a schematic view of an injection nozzle assembly
including a combustion dynamics reduction system in accordance with
yet another exemplary aspect of the invention.
[0014] The detailed description explains embodiments of the
invention, together with advantages and features, by way of example
with reference to the drawings.
DETAILED DESCRIPTION OF THE INVENTION
[0015] The terms "axial" and "axially" as used in this application
refer to directions and orientations extending substantially
parallel to a center longitudinal axis of a centerbody of a burner
tube assembly. The terms "radial" and "radially" as used in this
application refer to directions and orientations extending
substantially orthogonally to the center longitudinal axis of the
centerbody. The terms "upstream" and "downstream" as used in this
application refer to directions and orientations relative to an
axial flow direction with respect to the center longitudinal axis
of the centerbody.
[0016] With initial reference to FIG. 1, a turbomachine constructed
in accordance with exemplary embodiments of the invention is
generally indicated at 2. Turbomachine 2 includes a compressor 4
and a combustor assembly 5 having at least one combustor 6 provided
with an injection nozzle assembly housing 8. Turbomachine 2 also
includes a turbine 10 and a common compressor/turbine shaft 12. In
one exemplary embodiment, turbomachine 2 is a PG9371 9FBA Heavy
Duty Gas Turbine Engine, commercially available from General
Electric Company, Greenville, S.C. Notably, the present invention
is not limited to any one particular engine and may be used in
connection with other turbomachines.
[0017] As best shown in FIG. 2, combustor 6 is coupled in flow
communication with compressor 4 and turbine 10. Compressor 4
includes a diffuser 22 and a compressor discharge plenum 24 that
are coupled in flow communication with each other. Combustor 6 also
includes an end cover 30 positioned at a first end thereof, and a
cap member 34. Combustor 6 further includes a plurality of
pre-mixers or injection nozzle assemblies, two of which are
indicated at 38 and 39. Each injection nozzle assembly 38, 39
includes a corresponding main body 40, 41 having first and second
end portions 42, 43 and 44, 45 respectively. Second end portions 44
and 45 define an exit plane (not separately labeled) of injection
nozzle assemblies 38 and 39 respectively. In addition, combustor 6
includes a combustor casing 46 and a combustor liner 47. As shown,
combustor liner 47 is positioned radially inward from combustor
casing 46 so as to define a combustion chamber 48. An annular
combustion chamber cooling passage 49 is defined between combustor
casing 46 and combustor liner 47. Combustor 6 is coupled to
turbomachine 2 through a transition piece 55. Transition piece 55
channels combustion gases generated in combustion chamber 48
downstream towards a first stage turbine nozzle 62. Towards that
end, transition piece 55 includes an inner wall 64 and an outer
wall 65. Outer wall 65 includes a plurality of openings 66 that
lead to an annular passage 68 defined between inner wall 64 and
outer wall 65. Inner wall 64 defines a guide cavity 72 that extends
between combustion chamber 48 and turbine 10.
[0018] During operation, air flows through compressor 4, is
compressed, and passed to combustor 6 and, more specifically, to
injector assemblies 38 and 39. At the same time, fuel is passed to
injector assemblies 38 and 39 to mix with the compressed air to
form a combustible mixture. The combustible mixture is channeled to
combustion chamber 48 and ignited to form combustion gases. The
combustion gases are then channeled to turbine 10. Thermal energy
from the combustion gases is converted to mechanical rotational
energy that is employed to drive compressor/turbine shaft 12.
[0019] More specifically, turbine 10 drives compressor 4 via
compressor/turbine shaft 12 (shown in FIG. 1). As compressor 4
rotates, compressed air is discharged into diffuser 22 as indicated
by associated arrows. In the exemplary embodiment, a majority of
the compressed air discharged from compressor 4 is channeled
through compressor discharge plenum 24 towards combustor 6. Any
remaining compressed air is channeled for use in cooling engine
components. Compressed air within discharge plenum 24 is channeled
into transition piece 55 via outer wall openings 66 and into
annular passage 68. The compressed air is then channeled from
annular passage 68 through annular combustion chamber cooling
passage 49 and to injection nozzle assemblies 38 and 39. The fuel
and air are mixed to form the combustible mixture. The combustible
mixture is ignited to form combustion gases within combustion
chamber 48. Combustor casing 47 facilitates shielding combustion
chamber 48 and its associated combustion processes from the outside
environment such as, for example, surrounding turbine components.
The combustion gases are channeled from combustion chamber 48
through guide cavity 72 and towards turbine nozzle 62. The hot
gases impacting first stage turbine nozzle 62 create a rotational
force that ultimately produces work from turbomachine 2. At this
point it should be understood that the above-described construction
is presented for a more complete understanding of exemplary
embodiments of the invention.
[0020] As best shown in FIG. 3, turbomachine 2 includes a
combustion dynamics reduction system 90. In accordance with one
exemplary embodiment, combustion dynamics reduction system 90
includes a boundary layer perturbation mechanism 96 shown in the
form of an air/inert injection system 100. More specifically,
injection nozzle assembly 38 includes an outer conduit 104 and an
inner conduit 106 between which is defined a passage 108 having an
inlet (not separately labeled). Passage 108 has an outlet or
opening 112 arranged at second end portion 44 of injection nozzle
assembly 38. With this arrangement, air/inert is injected into
passage 108 and guided toward second end portion 44. The air/inert
passes through opening 112 towards the exit plane of injection
nozzle assembly 38. The air creates a disruption or disturbance at
a boundary layer between a flame present at the exit plane and the
unignited combustible mixture. More specifically, the air/inert
disrupts flow patterns of the fuel-air mixture within injection
nozzle assembly 38. By controlling air/inert flow rate through
passage 108, air/inert injection system 100 alters out vortex
shedding behavior at the boundary layer present at the base portion
of the flame. The disruption of the boundary layer and vortex
characteristics de-couples a combustion field component and an
acoustic component present at second end portion 44 in order to
suppress any attendant combustion screech.
[0021] Reference will now be made to FIG. 4 in describing a
boundary layer perturbation mechanism 120 in accordance with
another exemplary embodiment. As shown, boundary layer perturbation
system 120 includes a plurality of mechanical members 122 arranged
at second end portion 45 of injection nozzle assembly 39. In a
manner similar to that described above, injection nozzle assembly
39 includes an outer conduit 130 having an inner passage 131 that
leads to a discharge portion 132 provided at second end portion 45.
Injection nozzle assembly 39 also includes an inner conduit 135
having an internal passage 136 that leads to a discharge section
137 also arranged adjacent second end portion 45. With this
arrangement, the plurality of mechanical members 122 are arranged
on inner surfaces (not separately labeled) of outer conduit 130 as
well as outer surfaces (not separately labeled) of inner conduit
135. In accordance with one aspect of the exemplary embodiment, the
plurality of mechanical members take the form of protrusions 142.
However, mechanical members 122 could also take the form of
turbulators that trip the boundary layer into turbulance and/or
flappers that impart a pulsation motion to the fuel/air mixture and
may also disrupt the boundary layer. In this manner, a perturbation
effect is imparted to a fuel/air mixture passing through injection
nozzle assembly 39 prior to being released into combustion chamber
48 and ignited. The perturbation effect within injection nozzle
assembly 39 and disruption/alteration of boundary layer and
associated vortex structure in combustor 48 results in a decoupling
of combustion and acoustic components of the combustion process in
order to suppress combustion screech in turbomachine 2.
[0022] Reference will now be made to FIG. 5, wherein like reference
numbers represent corresponding parts in the separate views, in
describing a combustion dynamics reduction system 164 constructed
in accordance with yet another embodiment of the invention. As
shown, combustion dynamics reduction system 164 includes an
acoustic wave introduction system 167 and a fluid introduction
system 169. More specifically, acoustic wave introduction system
167 includes a first input line 177 having a first end 180 that
extends to a second end 181. Second end 181 is fluidly connected to
first end portion 42 of injection nozzle assembly 38. A valve 182
is positioned within first input line 177 to control introduction
of an acoustic wave that is delivered into injection nozzle
assembly 38. More specifically, acoustic wave introduction system
167 includes an acoustic driver 185 that is positioned at first end
180 of first input line 177. In a manner that will be described
more fully below, acoustic driver 185 is selectively operated to
deliver acoustic waves at various frequencies to a base of the
flame present at injection nozzle assembly 38.
[0023] As further shown in FIG. 5, fluid introduction system 169
includes a second input line 190 having a first end 193 that
extends to a second end 194. In a manner similar to that described
above, second end 194 is fluidly connected to first end portion 42
of injection nozzle assembly 38. Second input line 190 includes a
valve 195 that controls the introduction of a fluid, such as air,
fuel, and/or diluents, into injection nozzle assembly 38. With this
arrangement an acoustic wave and/or air/fuel/diluent or mixture
thereof is introduced into injection nozzle assembly 38 and
directed at a base portion of the flame within combustion chamber
48. The introduction of the acoustic wave and/or fluid decouples
the acoustic component and the combustion field component in order
to suppress combustion screech within turbomachine 2. More
specifically, acoustic driver 185 is operated to change both
frequency and amplitude of an acoustic wave passing into injection
nozzle assembly 38 in order to perturb or disrupt the base of the
flame within combustion chamber 48. Likewise, air can also be
injected into injection nozzle assembly 38 to further impact the
base of the flame and associated boundary layer and vortex
characteristics. This alters the response of the combustion
component and de-couples the combustion component from the acoustic
component.
[0024] At this point it should be understood that the present
invention provides a system for suppressing combustion screech in a
turbomachine by creating a boundary layer disruption within
injection nozzles associated with a particular combustor or
providing a system to disrupt a base portion of the flame directly
at an exit of a particular injection nozzle. By creating time
varying changes within the injection nozzle assemblies, combustion
screech can be significantly reduced if not eliminated.
Furthermore, eliminating combustion screech in this manner allows
operators to take advantage of all turbomachine operating ranges.
In addition, by suppressing combustion screech at the source, i.e.
within the nozzle assembly and/or combustion chamber, and
development of a high frequency dynamic field is eliminated before
having a chance to propagate through turbomachine components.
[0025] While the invention has been described in detail in
connection with only a limited number of embodiments, it should be
readily understood that the invention is not limited to such
disclosed embodiments. Rather, the invention can be modified to
incorporate any number of variations, alterations, substitutions or
equivalent arrangements not heretofore described, but which are
commensurate with the spirit and scope of the invention.
Additionally, while various embodiments of the invention have been
described, it is to be understood that aspects of the invention may
include only some of the described embodiments. Accordingly, the
invention is not to be seen as limited by the foregoing
description, but is only limited by the scope of the appended
claims.
* * * * *