U.S. patent application number 14/038056 was filed with the patent office on 2014-04-03 for variable flow divider mechanism for a multi-stage combustor.
This patent application is currently assigned to PETER JOHN STUTTAFORD. The applicant listed for this patent is YAN CHEN, STEPHEN JORGENSEN, PETER JOHN STUTTAFORD. Invention is credited to YAN CHEN, STEPHEN JORGENSEN, PETER JOHN STUTTAFORD.
Application Number | 20140090400 14/038056 |
Document ID | / |
Family ID | 50383939 |
Filed Date | 2014-04-03 |
United States Patent
Application |
20140090400 |
Kind Code |
A1 |
STUTTAFORD; PETER JOHN ; et
al. |
April 3, 2014 |
VARIABLE FLOW DIVIDER MECHANISM FOR A MULTI-STAGE COMBUSTOR
Abstract
The present invention discloses a novel apparatus and way for
altering the airflow to a gas turbine combustion system. The
apparatus comprises a flow divider mechanism which splits the
airflow surrounding a combustion liner into two distinct portions,
one directed towards a pilot and one directed towards a main stage
combustion. The flow divider mechanism is interchangeable so as to
provide a way of altering airflow splits between stages of the
combustion system.
Inventors: |
STUTTAFORD; PETER JOHN;
(JUPITER, FL) ; JORGENSEN; STEPHEN; (PALM CITY,
FL) ; CHEN; YAN; (WOODINVILLE, WA) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
STUTTAFORD; PETER JOHN
JORGENSEN; STEPHEN
CHEN; YAN |
JUPITER
PALM CITY
WOODINVILLE |
FL
FL
WA |
US
US
US |
|
|
Assignee: |
STUTTAFORD; PETER JOHN
JUPITER
FL
CHEN; YAN
WOODINVILLE
WA
JORGENSEN; STEPHEN
PALM CITY
FL
|
Family ID: |
50383939 |
Appl. No.: |
14/038056 |
Filed: |
September 26, 2013 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
61708323 |
Oct 1, 2012 |
|
|
|
Current U.S.
Class: |
60/796 ;
60/751 |
Current CPC
Class: |
F23R 3/34 20130101; F23C
2201/20 20130101; F23R 3/286 20130101; F23R 3/343 20130101; F23R
3/14 20130101; F23C 2900/07001 20130101; F23C 2900/06043 20130101;
F23R 3/54 20130101; F23R 2900/00014 20130101; F23R 3/60 20130101;
F23R 3/16 20130101; F23R 2900/03343 20130101; F23R 3/26
20130101 |
Class at
Publication: |
60/796 ;
60/751 |
International
Class: |
F23R 3/26 20060101
F23R003/26; F23R 3/60 20060101 F23R003/60 |
Claims
1. A flow divider mechanism comprising an annular plate positioned
about a combustion liner for dividing airflow into a pilot stage
and a main combustion stage of a gas turbine combustor, the annular
plate having a central opening, an outer edge, a first plurality of
openings located about the central opening, a second plurality of
openings located radially outward of the first plurality of
openings, and a third plurality of openings located adjacent the
outer edge, wherein the first plurality of openings and second
plurality of openings are sized to regulate and direct a
predetermined amount of airflow through multiple stages of the gas
turbine combustor.
2. The flow divider mechanism of claim 1, wherein the second
plurality of openings are offset circumferentially from the first
plurality of openings.
3. The flow divider mechanism of claim 1, wherein compressed air
for use in generating a main stage combustion flame passes through
the first plurality of openings in the annular plate.
4. The flow divider mechanism of claim 3, wherein compressed air
for use in generating and supporting a pilot flame passes through
the second plurality of openings in the annular plate.
5. The flow divider mechanism of claim 1 further comprising a flow
separator extending co-annular with the annular plate and
perpendicular relative to the annular plate.
6. The flow divider mechanism of claim 1, wherein the third
plurality of openings are used for clocking and securing the flow
divider mechanism to the gas turbine combustor.
7. The flow divider mechanism of claim 1, wherein the first
plurality of openings are in alignment with a corresponding main
stage mixing vane.
8. A multi-stage combustion system for directing a predetermined
amount of compressed air from outside of a combustion liner to
multiple stages within the combustion liner, the combustion system
comprising: a flow sleeve surrounding the combustion liner; a flow
divider mechanism positioned axially between the flow sleeve and a
main injector, the flow divider mechanism comprising an annular
plate positioned about the combustion liner for dividing airflow
passing between the flow sleeve and the combustion liner into a
first portion and a second portion, the annular plate having a
central opening, an outer edge, a first plurality of openings
located about the central opening, a second plurality of openings
located radially outward of the first plurality of openings, a
third plurality of openings located adjacent an outer edge; and a
cylindrical flow separator extending from the annular plate and
towards an inlet end of the combustion liner; wherein compressed
air passing between an outer wall of the combustion liner and the
flow sleeve is split into two portions, with a first portion
directed through the first plurality of openings and a second
portion directed through the second plurality of openings, the
first portion supplying compressed air to a main stage of
combustion and the second portion supplying air to a pilot
stage.
9. The combustion system of claim 8 further comprising a dome
having a hemispherical portion which causes a reversal in flow
direction of the first portion of compressed air.
10. The combustion system of claim 9, wherein the first portion of
compressed air passes along an outer wall of the combustion liner
when external to the combustion liner and along an inner wall of
the combustion liner after encountering the dome.
11. The combustion system of claim 10, wherein the second portion
of compressed air passes radially outward of the first portion of
compressed air when external to the combustion liner and radially
inward of the first portion of compressed air when internal of the
combustion liner.
12. The combustion system of claim 8, wherein the first plurality
of openings are in airflow alignment with a corresponding main
stage mixing vane.
13. The combustion system of claim 8, wherein the flow divider
mechanism is secured to the combustion system using the third
plurality of openings.
14. The combustion system of claim 13, wherein the flow divider
mechanism is interchangeable upon disengagement of surrounding
combustion hardware and fasteners securing the flow divider
mechanism to the combustion system.
15. A method of altering an airflow distribution between multiple
stages of a combustion system comprising: providing a combustion
system having a first flow divider mechanism in which compressed
air for use in combustion is divided into a first portion and a
second portion by an annular plate having a first plurality of
openings and a second plurality of openings; removing a cover,
dome, main fuel injector, and pilot nozzle from the combustion
system; removing fasteners securing the first flow divider to the
combustion system; removing the first flow divider; placing a
second flow divider on the combustion system, the second flow
divider having a first plurality of openings and a second plurality
of openings, where at least one of the first plurality of openings
or the second plurality of the second flow divider differs from the
first plurality of openings or a second plurality of openings of
the first flow divider; securing the second flow divider to the
combustion system; and securing the cover, dome, main fuel injector
and pilot nozzle to the combustion system such that the second flow
divider is positioned axially between flanges of the main fuel
injector and a flow sleeve.
16. The method of claim 15, wherein the second plurality of
openings in the second flow divider has an effective flow area
greater than an effective flow area for the second plurality of
openings in the first flow divider.
17. The method of claim 15, wherein the second plurality of
openings in the second flow divider has an effective flow area less
than an effective flow area for the second plurality of openings in
the first flow divider.
18. The method of claim 15, wherein the first plurality of openings
in the second flow divider has an effective flow area greater than
an effective flow area for the first plurality of openings in the
first flow divider.
19. The method of claim 15, wherein the first plurality of openings
in the second flow divider has an effective flow area less than an
effective flow area for the first plurality of openings in the
first flow divider.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This application claims the benefit of U.S. Provisional
Patent Application Ser. No. 61/708,323 filed on Oct. 1, 2012.
TECHNICAL FIELD
[0002] The present invention relates generally to an apparatus and
method for directing a predetermined airflow into a multi-stage gas
turbine combustion system. More specifically, an interchangeable
plate is positioned within the air flow path, external of the
combustion process, to split the air flow between a main combustor
stage and a pilot stage.
BACKGROUND OF THE INVENTION
[0003] In an effort to reduce the amount of pollution emissions
from gas-powered turbines, governmental agencies have enacted
numerous regulations requiring reductions in the amount of oxides
of nitrogen (NOx) and carbon monoxide (CO). Lower combustion
emissions can often be attributed to a more efficient combustion
process, with specific regard to fuel injector location, airflow
rates, and mixing effectiveness.
[0004] Early combustion systems utilized diffusion type nozzles,
where fuel is mixed with air external to the fuel nozzle by
diffusion, proximate the flame zone. Diffusion type nozzles produce
high emissions due to the fact that the fuel and air burn
essentially upon interaction, without mixing, and
stoichiometrically at high temperature to maintain adequate
combustor stability and low combustion dynamics.
[0005] An enhancement in combustion technology is the concept of
premixing fuel and air prior to combustion to form a homogeneous
mixture that burns at a lower temperature than a diffusion type
flame and thereby produces lower NOx emissions. Premixing can occur
either internal to the fuel nozzle or external thereto, as long as
it is upstream of the combustion zone. An example of a premixing
combustor of the prior art is shown in FIG. 1. A combustor 100 has
a plurality of fuel nozzles 102, each injecting fuel into a premix
cavity 104 where fuel mixes with compressed air 106 from plenum 108
before entering combustion chamber 110. Premixing fuel and air
together before combustion allows for the fuel and air to form a
more homogeneous mixture, which, when ignited will burn more
completely, resulting in lower emissions. However, in this
configuration the fuel is injected in relatively the same plane of
the combustor, and prevents any possibility of improvement through
altering the mixing length.
[0006] An alternate means of premixing fuel and air and obtaining
lower emissions can occur by utilizing multiple combustion stages.
In order to provide a combustor with multiple stages of combustion,
the fuel and air, which mix and burn to form the hot combustion
gases, must also be staged. By controlling the amount of fuel and
air passing into the combustion system, available power as well as
emissions can be controlled. Fuel can be staged through a series of
valves within the fuel system or dedicated fuel circuits to
specific fuel injectors. Air, however, can be more difficult to
stage given the large volume of air supplied by the engine
compressor. In fact, because of the general design to gas turbine
combustion systems, as shown by FIG. 1, air flow to a combustor is
typically controlled by the size of the openings in the combustion
liner itself, and is therefore not readily adjustable.
SUMMARY
[0007] The present invention discloses an apparatus and method for
controlling the amount of airflow directed into a multi-stage
combustion system. More specifically, in an embodiment of the
present invention, a flow divider mechanism is provided comprising
an annular plate positioned about a combustion liner having a first
plurality of openings for regulating airflow to a main stage of the
combustion system while a second plurality of openings are located
radially outward of the first plurality of openings and regulate
airflow to a pilot stage of the combustion system. The flow divider
mechanism is secured to the gas turbine combustion system in a way
such that it is removable and can be replaced in the field thereby
changing the airflow distribution to the combustion system.
[0008] In an alternate embodiment of the present invention, a
multi-stage combustion system is provided in which airflow to
multiple stages of the combustion system is regulated outside of a
combustion liner. The combustion system comprises a flow sleeve
surrounding a combustion liner and a flow divider mechanism for
directing airflow into a pilot stage and a main combustion stage
and a cylindrical flow separator extending from the flow divider
mechanism towards an inlet of the combustion liner.
[0009] In yet another embodiment of the present invention, a method
of altering an airflow distribution between multiple stages of a
combustion system is disclosed. The method comprises providing a
combustion system having a first flow divider mechanism capable of
dividing airflow between two stages of a combustor, removing a
portion of the combustion system in order to access the first flow
divider mechanism, removing the first flow divider mechanism and
replacing it with a second flow divider mechanism having different
airflow characteristics than the first flow divider mechanism. The
portion of the combustion system that was removed is then
reinstalled and the engine is returned to operation.
[0010] Additional advantages and features of the present invention
will be set forth in part in a description which follows, and in
part will become apparent to those skilled in the art upon
examination of the following, or may be learned from practice of
the invention. The instant invention will now be described with
particular reference to the accompanying drawings.
BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS
[0011] The present invention is described in detail below with
reference to the attached drawing figures, wherein:
[0012] FIG. 1 is a cross section of a portion of a gas turbine
engine and combustion system of the prior art.
[0013] FIG. 2 is a cross section of a gas turbine combustor in
accordance with an embodiment of the present invention.
[0014] FIG. 3 is a cross section of a gas turbine combustor
depicting the multiple stages of operation for the combustor of
FIG. 2 in accordance with an embodiment of the present
invention.
[0015] FIG. 4 is a perspective view of a portion of the gas turbine
combustor of FIG. 2 in accordance with an embodiment of the present
invention.
[0016] FIG. 5 is a detailed cross section of a portion of the gas
turbine combustor of FIG. 2 in accordance with an embodiment of the
present invention.
[0017] FIG. 6 is a cross section view of the gas turbine combustor
of FIG. 4 in accordance with an embodiment of the present
invention.
[0018] FIG. 7 is an end view of a flow divider mechanism in
accordance with an embodiment of the present invention.
[0019] FIG. 8 is a partial cross section view of the variable flow
meterplate of FIG. 7 in accordance with an embodiment of the
present invention.
[0020] FIG. 9 is a flow chart depicting a process of by which
airflow to the combustion system is changed in accordance with an
embodiment of the present invention.
DETAILED DESCRIPTION
[0021] By way of reference, this application incorporates the
subject matter of U.S. Pat. Nos. 6,935,116, 6,986,254, 7,137,256,
7,237,384, 7,308,793, 7,513,115, and 7,677,025.
[0022] The present invention discloses an apparatus for and way of
regulating and adjusting the airflow distribution to multiple
stages of a gas turbine combustion system. That is, embodiments of
the invention disclosed provide means for distributing the airflow
to stages of the combustor and altering the airflow to the
combustion system when it is determined airflow levels to one or
more stages of the combustion system should change.
[0023] The present invention will now be discussed with respect to
FIGS. 2-8. An embodiment of a gas turbine combustion system 200 on
which the present invention operates is depicted in FIG. 2. The
combustion system 200 is an example of a multi-stage combustion
system. The combustion system 200 extends about a longitudinal axis
A-A and includes a flow sleeve 202 for directing a predetermined
amount of compressor air along an outer surface of a combustion
liner 204. Compressor air then passes through a flow divider
mechanism 206 before a portion of the air mixes with fuel from main
fuel injectors 208. The flow divider mechanism 206 is discussed in
greater detail below. The divided portions of the flow exiting the
airflow divider mechanism 206 remain divided due to a generally
cylindrical flow separator 210 that extends from the flow divider
mechanism 206 and forward towards an inlet end 212 of the
combustion liner 204.
[0024] The combustion system 200 also comprises a dome 214 that is
positioned proximate the inlet end 212 of the combustion liner 204.
The dome 214 has a hemispherical cross-sectional shape such that
when encountered by a portion of the airflow, it causes the airflow
to reverse direction and enter the combustion liner 204.
[0025] The combustion system 200 also comprises a radially staged
premixer 216 with an end cover 218 having a first fuel plenum 220
extending about the longitudinal axis A-A of the combustion system
200 and a second fuel plenum 222 positioned radially outward of the
first fuel plenum 220 and concentric with the first fuel plenum
220. The radially staged premixer 216 also comprises a radial
inflow swirler 224 having a plurality of vanes 226.
[0026] Extending generally along the longitudinal axis A-A is a
pilot fuel nozzle 228 for providing and maintaining a pilot flame
for the combustion system. The pilot flame is used to ignite,
support and maintain the main combustion flame generated by
multiple stages from main fuel injectors 208.
[0027] As one skilled in the art understands, a gas turbine engine
typically incorporates a plurality of combustors. Generally, for
the purpose of discussion, the gas turbine engine may include low
emission combustors such as those disclosed herein and may be
arranged in a can-annular configuration about the gas turbine
engine. One type of gas turbine engine (e.g., heavy duty gas
turbine engines) may be typically provided with, but not limited
to, six to eighteen individual combustors, each of them fitted with
the components outlined above. Accordingly, based on the type of
gas turbine engine, there may be several different fuel circuits
utilized for operating the gas turbine engine. The combustion
system 200 disclosed in FIGS. 2 and 3 is a multi-stage premixing
combustion system comprising four stages of fuel injection based on
the loading of the engine. However, it is envisioned that the
specific fuel circuitry and associated control mechanisms could be
modified to include fewer or additional fuel circuits.
[0028] The pilot fuel nozzle 228 is connected to a fuel supply (not
shown) and provides fuel to the combustion system 200 for supplying
a pilot flame 250 where the pilot flame 250 is positioned generally
along the longitudinal axis A-A. The radially staged premixer 216
including the fuel plenums 220 and 222, radial inflow swirler 224
and its plurality of vanes 226 provide a fuel-air mixture through
the vanes 226 for supplying additional fuel to the pilot flame 250
by way of a pilot tune stage, or P-tune, 252.
[0029] As discussed above, combustion system 200 also includes main
fuel injectors 208. For the embodiment of the present invention
shown in FIG. 2, the main fuel injectors 208 are located radially
outward of the combustion liner 204 and spread in an annular array
about the combustion liner 204. The main fuel injectors 208 are
divided into two stages with a first stage extending approximately
120 degrees about the combustion liner 204 and a second stage
extending the remaining annular portion, or approximately 240
degrees, about the combustion liner 204. The first stage of the
main fuel injectors 208 are used to generate a Main 1 flame 254
while the second stage of the main fuel injectors 208 generate a
Main 2 flame 256.
[0030] As discussed above, the present invention provides a flow
divider mechanism 206 for regulating and splitting the amount of
compressed air supplied to different parts of the combustion liner
204. A flow divider mechanism 206, in accordance with an embodiment
of the present invention, is shown in detail in FIGS. 4 and 6-8.
The flow divider mechanism 206 comprises an annular plate 230
positioned about the combustion liner 204 and configured to divide
a passing airflow between the pilot stage 250/pilot tune stage 252
and the Main 1 and Main 2 combustion stages, 254 and 256,
respectively. For the embodiment of the present invention shown in
FIGS. 4 and 6-8, the annular plate 230 has a central opening 232,
an outer edge 234, and a first plurality of openings 236 that are
located about the central opening 232. As it can be seen from FIG.
7, the first plurality of openings 236 have a generally rectangular
cross section and extend radially outward from adjacent the central
opening 232. Although the first plurality of openings can be of
different shapes, a radially oriented generally rectangular cross
sectional opening maximizes the available flow area for the
material of the annular plate 230. Furthermore, for the embodiment
of the present invention shown in FIGS. 4 and 6-8, the first
plurality of openings 236, through which compressed air for use in
generating a main combustion flame (Main 1 and/or Main 2) passes,
are preferably in alignment with a corresponding main stage mixing
vane (not shown).
[0031] Referring back to FIG. 7, the annular plate 230 further
comprises a second plurality of openings 238 located radially
outward of the first plurality of openings 236. The second
plurality of openings 238 regulate the amount of cooling air that
is being passed into a passage supplying air to and in support of
the pilot flame 250 and pilot tune stage 252. The second plurality
of openings 238 may have a generally rectangular or circular cross
section oriented so as to extend radially outward. For the
embodiment of the annular plate 230 depicted in FIG. 7, the second
plurality of openings 238 are offset circumferentially from the
first plurality of openings 236, but the first and second plurality
of openings can also be in radial alignment. However, as discussed
above with respect to first plurality of openings 236, the second
plurality of openings 238 can also vary in size and shape,
depending on the airflow requirements and available area in the
annular plate 230.
[0032] The configuration of the annular plate 230 is generally a
flat plate having a nominal thickness that should be accounted for
in determining flow split. The present invention provides a means
for thickness to be accounted for as a varying parameter in the
design phase, and as such, is not limited to a specific thickness
range.
[0033] The size and shape of the first plurality of openings 236
and the second plurality of openings 238 depends on a variety of
conditions, such as size of the combustion system, desired fuel-air
mixing levels, and required airflow to various stages of the
combustion system, among others. Therefore, the shape of the
openings 236 and 238 and their corresponding effective flow area
will vary. In one embodiment, it is envisioned that approximately
60% of the compressed air passing through the flow divider
mechanism 206 is directed through the first plurality of openings
236 with the remaining approximately 40% of compressed air directed
through the second plurality of openings 238. In alternate
embodiments of the present invention, fewer or more openings can be
located in the annular plate than those shown in the enclosed
figures, such as arc-shaped openings to further increase the
effective flow area.
[0034] As discussed above and referring back to FIG. 2, the airflow
exits the flow divider mechanism 206 in divided portions. The
airflow portions remain separated due to the generally cylindrical
flow separator 210 that extends from the flow divider mechanism 206
and forward towards an inlet end 212 of the combustion liner
204.
[0035] Referring back to FIG. 7, the annular plate 230 of the flow
divider mechanism 206 further comprises a third plurality of
openings 240 located adjacent the outer edge 234. Instead of
regulating airflow, the third plurality of openings 240 are used
for properly orienting and securing the flow divider mechanism 206
on the combustion system 200. The flow divider mechanism 206 is
secured to the combustion system 200 by a plurality of removable
fasteners (not shown).
[0036] As it can be seen from FIGS. 2 and 5, the flow divider
mechanism 206 is positioned axially between a flange of the flow
sleeve 202 and the main injector 208 such that the annular plate
230 of the flow divider mechanism 206 is essentially sandwiched
between adjacent components of the combustion system 200. The
fasteners 207 for securing the flow divider mechanism 206 pass
through the third plurality of openings 240 and engage openings in
the flow sleeve 202.
[0037] As mentioned briefly above, the combustion system 200
includes a dome 214 having a hemispherical shape. The dome 214
provides a means for reversing a portion of the airflow passing
through the flow divider mechanism 206. More specifically, the
first portion of air, which passes through the first plurality of
openings 236 initially passes along an outer wall 204A of the
combustion liner 204 while external to the combustion liner and
then, due to the dome 214, reverses direction and passes along an
inner wall 204B of the combustion liner 204. The portion of the
compressed air passing through the second plurality of openings 238
initially passes radially outward of the first portion of the
compressed air when external to the combustion liner 204, but is
then positioned radially inward of this first portion of the
compressed air once inside the combustion liner 204. While the dome
214 is used to provide a flow reversal mechanism to the portion of
the compressed air passing through the first plurality of openings
236, the portion of the air which passes through the second
plurality of openings 238 reverses flow direction into the
combustion liner 204 as a result of passing through the radial
inflow swirler 224.
[0038] In addition to the ability to regulate the amount of
compressed air passing into each of the respective circuits of the
combustion system, the present invention also provides a way of
modifying or adjusting the airflow distribution between multiple
stages of a combustion system. Referring to FIG. 9, the process 900
for altering the airflow distribution to the combustion system 200
is provided. Initially, in a step 902, the combustion system having
a first flow divider mechanism is provided. This combustion system
and first flow divider mechanism is similar to that previously
described. Then, in a step 904, a determination is made that a
change to the airflow to the combustion system is required. This
determination may be made due to a variety of factors such as
emissions levels, combustion noise, and turndown, among others.
[0039] Once it has been determined that the airflow split between
the pilot and main combustion stages must be changed, in order to
access the flow divider mechanism, the cover, dome, main fuel
injector and pilot fuel nozzle are removed in a step 906. Once
these components have been removed, the flow divider mechanism is
accessible. Then, in a step 908, the fasteners securing the flow
divider mechanism to the combustion system are removed and in a
step 910, the first flow divider mechanism is removed.
[0040] In a step 912, a second flow divider mechanism is placed on
the combustion system. The second flow divider mechanism differs
from the first flow divider mechanism in that at least one of the
first plurality of openings and/or the second plurality of openings
in the second flow divider mechanism differ in size so as to alter
the overall effective flow area for the second flow divider
mechanism when compared to the first plurality of openings and/or
the second plurality of openings and effective flow area in the
first flow divider mechanism. Therefore, multiple combinations of
possible changes exist and can be made when switching from the
first flow divider mechanism to the second flow divider
mechanism.
[0041] In a step 914, the second flow divider mechanism is clocked
on the combustion system and secured to the combustion system using
fasteners, as discussed above. Once the second flow divider
mechanism has been secured to the combustion system, the cover,
dome, main fuel injector and pilot nozzle are secured to the
combustion system in a step 916.
[0042] Upon reinstallation of all combustion hardware, fuel lines
and any other hardware previously removed, the gas turbine engine
can be restarted using the existing controls programming. That is,
the changes to airflow to the combustion system are all hardware
changes such that little to no software changes should have to be
made with respect to the airflow changes. Slight changes in fuel
scheduling may be required in order to ensure emissions compliance
is maintained given the altered airflow configuration. If, upon
further operation and analysis, it is determined that there must be
another change to the airflow split of the combustion system, the
process outlined above can be repeated and the second flow divider
mechanism replaced with yet another flow divider mechanism.
[0043] While the invention has been described in what is known as
presently the preferred embodiment, it is to be understood that the
invention is not to be limited to the disclosed embodiment but, on
the contrary, is intended to cover various modifications and
equivalent arrangements within the scope of the following claims.
The present invention has been described in relation to particular
embodiments, which are intended in all respects to be illustrative
rather than restrictive.
[0044] From the foregoing, it will be seen that this invention is
one well adapted to attain all the ends and objects set forth
above, together with other advantages which are obvious and
inherent to the system and method. It will be understood that
certain features and sub-combinations are of utility and may be
employed without reference to other features and sub-combinations.
This is contemplated by and within the scope of the claims.
* * * * *