U.S. patent application number 14/857439 was filed with the patent office on 2016-07-14 for small arrayed swirler system for reduced emissions and noise.
The applicant listed for this patent is Rolls-Royce Canada, Ltd.. Invention is credited to Marc Furi, Sandeep Jella.
Application Number | 20160201918 14/857439 |
Document ID | / |
Family ID | 56367286 |
Filed Date | 2016-07-14 |
United States Patent
Application |
20160201918 |
Kind Code |
A1 |
Jella; Sandeep ; et
al. |
July 14, 2016 |
SMALL ARRAYED SWIRLER SYSTEM FOR REDUCED EMISSIONS AND NOISE
Abstract
A combustor assembly employs an ejector assembly having an array
of swirlers to influence the flow field, reduce unburned fuel
emissions, enhance flame stabilization and reduce noise within a
combustion chamber. The swirlers are located downstream of a
premixing section and are designed to correct bias of fuel-air
mixture residence times in the combustion chamber. The ejector
assembly may be tuned to any particular configuration by varying
the swirl number in order to produce recirculation zones that
ensure out of phase nature of the heat release within the
combustion chamber.
Inventors: |
Jella; Sandeep; (Montreal,
CA) ; Furi; Marc; (Dorval, CA) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
Rolls-Royce Canada, Ltd. |
Montreal |
|
CA |
|
|
Family ID: |
56367286 |
Appl. No.: |
14/857439 |
Filed: |
September 17, 2015 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
62052149 |
Sep 18, 2014 |
|
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Current U.S.
Class: |
60/737 |
Current CPC
Class: |
F23R 3/34 20130101; F23R
3/14 20130101; F23R 3/286 20130101; F23R 2900/00014 20130101 |
International
Class: |
F23R 3/28 20060101
F23R003/28; F23R 3/00 20060101 F23R003/00 |
Claims
1. A combustion apparatus comprising: a combustor liner; an
injector assembly including a primary combustion system and a
secondary combustion system; and a combustion chamber having an
axis extending therethrough.
2. The combustion apparatus according to claim 1, further
comprising a gas turbine engine.
3. The combustion apparatus according to claim 1, wherein the
primary combustion system includes a mixing chamber, and a nozzle,
the mixing chamber has ports for receiving air and fuel, the mixing
chamber further has an exit located adjacent to the nozzle.
4. The combustion apparatus according to claim 1, wherein the
primary combustion system includes an injector ring, the injector
ring is spaced apart from the combustor liner.
5. The combustion apparatus according to claim 1, wherein the
primary combustion system includes an injector member, the injector
member has an inside diameter, an outside diameter, a body, a front
face and a rear face, the front face has a plurality of passage
openings, the passage openings extend through the body.
6. The combustion apparatus according to claim 5, further
comprising a swirler located within at least one passage
opening.
7. The combustion apparatus according to claim 5, further
comprising a plurality of swirlers located around a circumference
of the body of the injector member.
8. The combustion apparatus according to claim 5, wherein the front
face is not normal to the axis that extends through the
combustor.
9. The combustion apparatus according to claim 5, further
comprising a swirler located at one portion of the injector member,
and at least one slot located at another portion of the injector
member.
10. The combustion apparatus according to claim 5, further
comprising an array of swirlers that are spaced apart from one
another around at least a portion of the injector member.
11. The combustion apparatus according to claim 5, wherein the
injector member is ring shaped and has an array of spaced apart
swirlers that are located downstream from a mixing channel.
12. The combustion apparatus according to claim 1, wherein the
secondary combustion system includes a mixing channel, a swirler,
and a passageway in communication with the mixing channel.
13. The combustion apparatus according to claim 1, wherein the
secondary combustion system has a flow path that delivers
combustible material to a secondary combustion zone, and the
primary combustion system has a flow path that delivers combustible
material to a primary combustion zone.
14. A combustion member comprising: a combustor liner; an injector
assembly, the injector assembly includes a body having a plurality
of spaced apart apertures circumferentially located around the
body, a swirler is located within at least one aperture; a primary
combustion system, the primary combustion system includes a mixing
space, the mixing space includes openings for receiving fuel and
air; and a secondary combustion system, the secondary combustion
system includes a mixing space that has openings for receiving fuel
and air.
15. The combustion apparatus according to claim 14, wherein the
injector assembly further includes an annular channel that is
spaced upstream of the swirler.
16. The combustion apparatus according to claim 14, further
including a swirler located within each aperture that is located
around the ring.
17. The combustion apparatus according to claim 14, wherein the
spaced apart apertures include a combination of slots and circular
shaped openings, the swirler is located within the circular shaped
opening, and the slot performs the function as a jet.
18. The combustion apparatus according to claim 14, wherein the
spaced apart further includes a swirler for introducing fuel
particles or fuel-air mixtures to a secondary combustion zone that
is located within a combustor.
19. The combustion apparatus according to claim 14, further
comprising a combustor shell.
20. A combustion member comprising: a combustor liner, the liner
has ports for facilitating the entry of air; an injector assembly,
the injector assembly includes an annular shaped ring having a
plurality of spaced apart apertures circumferentially located
around the ring, a swirler is located within each aperture; a
primary combustion system, the primary combustion system includes a
mixing space, the mixing space includes openings for receiving fuel
and air; a primary combustion zone located downstream of the
swirler; a secondary combustion system, the secondary combustion
system includes a second swirler and a mixing space that has
openings for receiving fuel and air; and a secondary combustion
zone located downstream of the second swirler.
Description
FIELD OF TECHNOLOGY
[0001] The present disclosure relates to gas turbine engines, and
more particularly, to an improved combustor assembly employing an
array of swirlers to influence the flow field and enhance flame
stabilization within a combustion chamber.
BACKGROUND
[0002] Gas turbine engines are known to include a compressor for
compressing air, a combustor for producing a hot gas by burning
fuel in the presence of the compressed air produced by the
compressor, and a turbine for expanding the hot gas to extract
power. Gas turbine engines using annular combustion systems
typically include a plurality of individual burners disposed in a
ring about an axial centerline for providing a mixture of fuel and
air to an annular combustion chamber disposed upstream of the
annular turbine inlet vanes. Other gas turbines use can-annular
combustors wherein individual burner cans feed hot combustion gas
into respective individual portions of the arc of the turbine inlet
vanes. Each can includes a plurality of main burners disposed in a
ring around a central pilot burner.
[0003] During operation, the combustion flame can generate
combustion oscillations, also known as combustion dynamics.
Combustion oscillations in general are acoustic oscillations which
are excited by the combustion itself. The frequency of the
combustion oscillations may be influenced by an interaction of the
combustion flame with the structure surrounding the combustion
flame. Since the structure of the combustor surrounding the
combustion flame is often complicated, and varies from one
combustor to another, and because the combustion flame itself may
vary over time, it is difficult to predict the frequency at which
combustion oscillations occur. As a result, combustion oscillations
may be monitored during operation and parameters may be adjusted in
order to influence the interaction of the combustion flame with its
environment.
[0004] A combustion flame emits sound energy during combustion. A
more uniform flame will generate more uniform acoustics, but
perhaps with higher peak amplitude at a particular frequency than a
less uniform flame. When an emitted frequency of combustion
coincides with a resonant frequency of the combustion chamber the
system may operate in resonance, and the resulting combustion
dynamics may damage the gas turbine components, or at least reduce
their lifespan.
[0005] One known way to reduce the interaction of the combustion
flame with the combustion acoustics is to reduce the coherence of
the flame, i.e. reduce the spatio-temporal uniformity of the flame.
A flame with less uniform combustion throughout its volume is
likely to perturb the gas turbine less than a uniform flame because
the energy released is spatially distributed and therefore
decreases its coupling to the system resonant frequencies or
acoustic modes. As a result, combustion dynamics of flames with
less uniform combustion throughout its volume are less likely to be
exacerbated than by a more uniform flame. Creating a less uniform
fuel-air mixture would be helpful.
[0006] During the combustion of gas, pollutants such as, but not
limited to, carbon monoxide ("CO.sub.2"), unburned hydrocarbons
("UHC"), and nitrogen oxides ("NO.sub.x") may be formed and emitted
into an ambient atmosphere. Because of stringent emission control
standards, it is desirable to control emissions of such pollutants
by the suppressing formation of such emissions. It would be helpful
to reduce such emissions to target levels so as to meet government
emission levels. This may be accomplished by minimizing the
resident time unburnt fuel resides within the combustion
chamber.
[0007] In the case of industrial gas turbines which primarily burn
gaseous fuels (e.g. for power generation purposes) several
contradictory requirements with respect to premixer, flame
stabilizer and combustion chamber design for lean
partially-premixed systems require an achievement of tradeoffs
between cost, complexity, robustness to operating conditions and
lowered emissions. For example, a perfectly premixed system is more
prone to thermo-acoustic oscillations leading to hardware damage
while a less uniformly mixed system results in increase of
pollutant formation, in particular, NOx. Another issue is that the
orientation of the combustion chamber within any engine
architecture can impose constraints on burner geometry, that result
in unsymmetrical stream tubes within the combustion chamber, which
in turn, lead to a variation of mixture residence times leading to
lowered combustion efficiencies, increased risk of undesirable
events such as poor flame stabilization, flashback, blow out and
less than adequate mixing of post-flame gases which is required for
the fast burnout of Carbon Monoxide to Carbon Dioxide. The time
taken for the latter is a hard constraint in many systems as it is
obtained--in a simple example--as the time taken for the mixture to
travel from the flame front to the chamber exhaust, which is
determined by the engine shaft length. However, overly long
residence times (desirable for oxidation of CO) can increase the
amount of NOx formation at high temperatures.
[0008] For this reason, it is helpful that the flame stabilizer to
include these considerations into its design and be able to
function in a robust manner in engine architectures of reasonable
variation. Typically, symmetric systems are the goal for reasons of
ease of operation and lowered costs but rarely achieved in
real-life. Still another tradeoff lies in ensuring that the mixture
in the premixer is expelled before it auto-ignites. At typical
engine operating conditions, the compressed air can reach over 800K
and pressures in excess of 500 psi. At these conditions, simple
fuels like CH4 can ignite within 40 milliseconds while more
reactive fuels such as diesel can ignite in less than 2
milliseconds. Premixer design should ensure that adequate
uniformity of mixing is achieved within a timescale that is lower
than the ignition delay of the range of fuels required. With safety
factors that can range from a tenth to twentieth of the timescales,
it is difficult to achieve a uniformity of mixture unless the
premixer design employs high turbulence intensities.
[0009] Fuel injection into a continuous burning combustion chamber
as, for example, in a gas turbine engine has posed continuing
design problems. Difficulties have been encountered in injecting
fuel in a highly dispersed manner so as to achieve complete and
efficient combustion of the fuel, and at the same time minimize the
occurrence of fuel-rich pockets which upon combustion produce
carbon, smoke, or unburned hydrocarbon pollutants. Fuel injection
difficulties have been further complicated by the introduction of
gas turbine engines having increased combustor pressure levels.
Existing fuel spray atomizer efficiency decreases as combustor
pressure is increased, resulting in a more non-uniform dispersion
of fuel, together with an increase in the fuel-rich zones within
the combustion chamber which cause reduced burner efficiency,
excessive exhaust smoke, and a non-uniform heating of the combustor
shell, a condition commonly referred to as hot streaking, which can
lead to rapid deterioration of the shell.
[0010] High fuel pressure spray atomizers also have not proved
entirely satisfactory because of the present limitations on fuel
pump pressure. Systems for vaporizing fuel upon injection into the
combustor have also proved to be severely limited due to the
dependence of the vaporization process on the temperature of the
fuel and air entering the combustor.
[0011] In view of the aforementioned challenges, there is a need to
provide an improved combustor assembly and method of operation that
provides improved flame stabilization for reducing emissions and
noise. The system may be tunable and include a plurality of
swirlers that results in a heat release that is designed to be out
of phase with each other so as to ensure that thermo-acoustic
couplings are dampened.
BRIEF DESCRIPTION OF THE DRAWINGS
[0012] While the claims are not limited to a specific illustration,
an appreciation of the various aspects is best gained through a
discussion of various examples thereof. Referring now to the
drawings, exemplary illustrations are shown in detail. Although the
drawings represent the illustrations, the drawings are not
necessarily to scale and certain features may be exaggerated to
better illustrate and explain an innovative aspect of an example.
Further, the exemplary illustrations described herein are not
intended to be exhaustive or otherwise limiting or restricted to
the precise form and configuration shown in the drawings and
disclosed in the following detailed description. Exemplary
illustrations are described in detail by referring to the drawings
as follows:
[0013] FIG. 1 illustrates a schematic view of a gas turbine engine
employing the improvements discussed herein;
[0014] FIG. 2 illustrates a partial perspective view of combustor
for a gas turbine engine employing an exemplary swirler assembly
for use with a combustor;
[0015] FIG. 3 illustrates a perspective view of a swirler ring
assembly having a plurality of swirlers disposed circumferentially
around a ring structure;
[0016] FIG. 4 is a side elevational view of the FIG. 3 swirler ring
assembly;
[0017] FIG. 5 is a front elevational view of the FIG. 3 swirler
ring assembly;
[0018] FIG. 6 is a front elevational view of an alternative swirler
ring assembly, employing an array of swirlers and slots;
[0019] FIG. 7 is a front elevational view of another alternative
swirler ring assembly, employing ports located adjacent to the
swirlers; and
[0020] FIG. 8 is a schematic view of a section of a combustor,
showing a primary section and a secondary section, and fluid flow
within the combustion chamber.
DETAILED DESCRIPTION
[0021] An exemplary embodiment discloses an improved combustor
assembly and method that overcomes traditional combustor challenges
by employing a tunable flame stabilizer assembly in the injector
section of an engine, such as, but not limited to, a gas turbine.
The stabilizer assembly includes an array of swirlers located
downstream of a premixing section that are designed to correct bias
of fuel residence times in the combustion chamber. The flame
stabilizer could employ an ejector like ring having a plurality of
circumferentially spaced swirlers that induce a premixed air/fuel
mixture to the combustion chamber. The stabilizer may be tuned by
varying the swirl number in order to produce recirculation zones
that ensure out of phase nature of the heat release within the
combustion chamber. Individual premixing channels receive fuel and
pre-heated air which rely on turbulent mixing set up by vortices
generated within the premixer channel and expels the mixed material
through each swirler and into the combustion chamber where it is
burned to provide energy for the turbine section.
[0022] An alternative embodiment provides a flame stabilizer
assembly featuring a hybrid arrangement of swirlers and jet nozzles
for passing the air/fuel mixture through an injector. This allows a
compensation of more drastic combustion chamber residence time
biases. Another embodiment features a premixer that injects
non-uniform fuel flow through a series of ports that are located
near the array of swirlers. It will be appreciated that a
combination of the embodiments disclosed herein may be employed so
as to include a variety of swirler, slot and injector
configurations so as to accommodate a wide variety of combustion
chambers typical of gas turbine engines
[0023] A method of operating an injector for a combustor includes
mixing air and fuel into a first primary mixing chamber,
introducing the air/fuel mixture into a plurality of swirlers,
directing the mixture into a combustion chamber, and then igniting
the mixture. The method further includes mixing air and fuel into a
secondary mixing chamber, introducing the mixed air and fuel into a
common secondary swirler, directing the mixture into a secondary
combustion zone within the combustion chamber, and igniting the
mixture. The method may further include additional primary mixing
chambers that are circumferentially spaced around the central axis
28 of the machine 10. Likewise, each primary mixing chamber has an
associated swirler that is associated with it which operates in a
manner similar to that discussed for the first primary mixing
chamber.
[0024] The resulting combustor assembly is an array of swirlers
that generates a series of small recirculation zones in order to
reduce the length of scales of coherent flame structures. The
number of swirlers can be varied in order to produce recirculation
zones that ensure out of phase nature of the heat released within
the combustor. The swirl number is tuned to the asymmetry of the
combustion chamber so that swirlers with a low swirl number are
situated at locations which give rise to stream tubes of high
residence time. By contrast, swirlers with a high swirl number are
situated at locations which rise to stream tubes of low residence
time. As such, the swirler may be used to correct bias of residence
times in the combustion chamber. It will be appreciated that the
number of swirlers may range from 0 to 18 (as shown), or more.
[0025] FIG. 1 illustrates a gas turbine engine 10, which includes a
fan 12, a low pressure compressor and a high pressure compressor,
14 and 16, a combustor 18, and a high pressure turbine and low
pressure turbine, 20 and 22, respectively. The high pressure
compressor 16 is connected to a first rotor shaft 24 while the low
pressure compressor 14 is connected to a second rotor shaft 26. The
shafts extend axially and are parallel to a longitudinal center
line axis 28.
[0026] Ambient air 30 enters the fan 12 and is directed across a
fan rotor 32 in an annular duct 34, which in part is circumscribed
by fan case 36. The bypass airflow 38 provides engine thrust while
the primary gas stream 40 is directed to the combustor 18 and the
high pressure turbine 20. The gas turbine engine 10 includes an
improved combustor 18 having a unique flame stabilizer assembly 42
for improved heat release in the combustion chamber.
[0027] FIG. 2 illustrates a perspective view of the inside of a
combustor assembly 18 showing some of the components of the flame
stabilizer assembly 42. It will be appreciated that the outer liner
wall has been removed so as to provide improved understanding of
the components of the stabilizer assembly 42.
[0028] The stabilizer assembly includes an injector ring 44, a
primary mixing chamber or duct 46, a secondary mixing chamber or
duct 48, a secondary swirler 50, a plurality of primary swirlers
52, and a liner 54. A primary combustion zone 56 connotes the area
in which the air/fuel mixture from an array primary swirlers 52
deposit atomized fuel particles or fuel-air mixtures that are ready
to be combusted. A secondary combustion zone 58 connotes an area
where the secondary swirler 50 deposits atomized fuel particles or
fuel-air mixtures that are ready to be combusted.
[0029] The secondary mixing chamber 48 receives air flow from air
duct 74 and fuel flow from fuel supply 64 in which the air and fuel
are mixed and exited at outlet 66. Vanes 68 force the mixture from
chamber 48 to exit through the outlet 66, thereby directing the
heated, pressurized fuel/air mixture or atomized fuel/air mixture
to be stabilized prior to combustion.
[0030] The inner liner 54 is made of conventional material and
extends between the secondary swirler 50 and the array of swirlers
52. The liner 54 extends axially from the injector ring 44 and
terminates near a flange 70 that is formed near one end of a
partition member 72. An air duct 74 supplies cooling air to the
inner liner 54 where ports 76 deliver the air to the primary
combustion zone 56. The secondary duct 48 receives secondary fuel
78 and secondary air 80 that collectively form the fuel/air mixture
64. See FIG. 8.
[0031] With reference to FIGS. 2 and 3, the injector ring 44 can
be, but is not limited to, an annularly shaped member that
circumscribes the inner liner 54. The ejector 44 has an inner
diameter 82 and an outer diameter 84 that are concentric with axis
28 and is uniform in width and depth but variations thereof to
accommodate non-circular jets are included as well. A plurality of
passages 86 extend from the face 88, through the body 90 of the
ejector 44, and extends to back 92 of the injector 44. Each passage
86 receives a swirler 52 for advancing the mixed air/fuel 64' that
is generated from the primary duct 46.
[0032] FIG. 4 illustrates a side view of the injector ring 44
whereby the face 88 has an angle alpha that forms a relief from
outer surface 94. This configuration differs from the flat face 88
that is shown in the embodiment of FIG. 2 whereby the flat face 88
is basically normal to the axis 28 of the machine. However, by
locating the face 88 at an angle alpha as is shown in FIG. 4, each
passage 86 is oriented to no longer be normal to the axis 28, but
instead is redirected to provide a path extending a flow of
atomized fuel or fuel-air mixture into the combustion chamber 18
that is non-normal. This causes small pockets of recirculation
zones that bleed over to adjacent recirculation zones from the
adjacent swirler 52. The angle alpha may be a variety of
configurations including, but not limited to 21, 24 or 26
degrees.
[0033] The injector ring 44 further has a first annular grove 96
and a second annular groove 98. The grooves extend around the
periphery of the ring 44 and form a fluid flow path. The ring 44 is
made of suitable material to withstand the temperatures that are
traditionally present in the combustor applications.
[0034] FIG. 5 illustrates the injector ring 44 from the front
elevational view. The ring 44 is shown in this exemplary embodiment
having 18 swirlers 52 that are spaced apart around the
circumference of the body 90. Each swirler is positioned within a
passage 86 for providing its own small recirculation zone within
the primary combustion zone 56. It will be appreciated that more or
less swirlers may be provided. While 5 blades 100 are shown on each
swirler 52, it will be appreciated that the number of blades 100 on
each swirler 52 may vary. Thus, the injector ring 44 may be tuned
by varying the number of swirlers 52 and blades 100 that are
employed. Also, the alpha angle may be varied so as to change the
pitch or trajectory 258 (FIG. 8) of delivery of atomized fuel air
mixture into the primary combustion zone 56.
[0035] FIG. 6 illustrates an alternative embodiment of an injector
ring 150 having a blend of swirlers 52 and slots 152. The swirler
designs are similar to those mentioned above, and in this
variation, 9 swirlers are presented in an array between the 6
o'clock to 12 o'clock positions. By contrast, an array of slots 152
are aligned between the 1 o'clock and 5 o'clock positions. It will
be appreciated that this mix of swirlers and slots may be
positioned at other locations and they could even be mixed
intermittently between one another. The slots 152 are shown as
arcuate shaped rectangles and they have a space 154 of material
separating each adjacent slot 152. The slots 152 are ports or jets
that extend through the body 90 and provide a passageway for the
jet flames to enter the primary combustion zone 56. The slots 152
may be fed with a fuel/air mixture at a different equivalence ratio
compared to the swirlers 52. The recirculation zone 202 (FIG. 8)
may be designed to have preferred performance characteristics that
are based upon a combination of swirlers 52 and slots 152.
[0036] FIG. 7 depicts another alternative embodiment injector ring
200 that is designed to inject a non-uniform fuel flow into the
primary combustor zone 56. This intentional mixture of
non-uniformity however is with small acceptable margins of
non-uniformity so as to contain the NOx emissions within industry
standards. This alternative embodiment 200 is a variant of the FIG.
5 assembly, whereby jets 204 and 206 are positioned adjacent to
swirlers 52 at positions around the circumference of the body 90.
The jets 204 and 206 are shown in pairs in alternating patterns
spaced adjacent to a pair of swirlers 52. It will be appreciated
that the jets may be positioned in between each swirler 52. The jet
204 is shown larger in diameter than jet 206 and they could be
reversed to have the smaller jet on the outside of the ring 200
with the larger jet 204 positioned near the inside diameter of the
ring 200.
[0037] FIG. 8 illustrates a schematic fluid diagram 250 of the
fuel/air mixtures entering the combustion zones within the
combustor 18. The primary mixer or duct 46 represents the fluid
mixing channel where primary fuel 252 and primary air 254 are mixed
to form the fuel/air mixture 64'. The mixed fuel 64' is delivered
to an injector ring of the styles shown at ejectors 44, 150, or
200. The mixture 64' is atomized in the process and passes through
the injector ring 44 to the primary combustion zone 56 which is at
a leading portion of the combustor 18.
[0038] The example shown has the injector outlet passage 86
configured at an angle alpha which results in the plume 256 of
atomized fuel or fuel-air mixture to enter the primary combustion
zone 56 along a trajectory 258. This results in small pockets 260
of recirculation zones that are associated with each swirler 52.
Thus, for each swirler 52 that is spaced around the periphery of
the injector ring 44, a small pocket 260 of recirculation zone is
induced into the primary combustion zone 56. The small pockets 260
overlap with one another so as to enhance combustion and to
maintain ignition within the combustor 18. The performance of each
small pocket 260 is controlled by, in part, the angle alpha, the
number and design of each blade for each swirler 52, and other
controllable features. As such the ejector ring 44 is tunable to
afford different performance characteristics which results in
different fluid flow patterns within the primary combustion zones
56 and the secondary combustion zone 58.
[0039] The secondary duct 48 is a mixing channel that receives the
secondary fuel 78 and the secondary air 80 which in turn is heated
and mixed to form a mixture 64 within the flow path 60. The mixture
64 is delivered to the jet or swirler 50 and exits into the
secondary combustion zone 58 and forms a recirculation pattern 262.
The recirculation patterns 262 and 202 may combine within the
trailing portion of the combustor 18 so as to enhance ignition of
the fuel particles and create heat for rotating the turbines.
[0040] The process of operating the engine 10 using the injector
ring 44 will now be presented with reference to schematic of FIG.
8. Fuel and air mixtures are induced to the primary duct 46 and
secondary duct 48. The fuel mixtures 64 and 64' advance to their
associated jets, 50 and 52, which in turn delivers atomized fuel
particles to their respective combustion zones 58 and 56. Ignition
now occurs within the combustion zones which produced heat for
driving the turbines 20 and 22.
[0041] The method of operation may be tuned by changing the
injector ring 44 to have a variety of swirlers 52 and slots 152
and/or jets 204 and 206, or any combination thereof. By modifying
this arrangement of components, the recirculation zone
characteristics 202 within the primary combustion zone 56 can be
tuned to a desired performance. This results in part due to the
individual recirculation zones 260 that are generated by each
individual swirler 52, slot 152 and or jets 204 and 206. Thus by
locating a swirler 52, slot 152 and/or jet 204, 206 at
predetermined locations around the ejector ring 44, the user can
influence the combustion, emission, and energy generated by the
engine 10.
[0042] It will be appreciated that the aforementioned method and
devices may be modified to have some components and steps removed,
or may have additional components and steps added, all of which are
deemed to be within the spirit of the present disclosure. Even
though the present disclosure has been described in detail with
reference to specific embodiments, it will be appreciated that the
various modifications and changes can be made to these embodiments
without departing from the scope of the present disclosure as set
forth in the claims. The specification and the drawings are to be
regarded as an illustrative thought instead of merely restrictive
thought.
[0043] All terms used in the claims are intended to be given their
broadest reasonable constructions and their ordinary meanings as
understood by those knowledgeable in the technologies described
herein unless an explicit indication to the contrary is made
herein. In particular, use of the singular articles such as "a,"
"the," "said," etc. should be read to recite one or more of the
indicated elements unless a claim recites an explicit limitation to
the contrary.
* * * * *