U.S. patent application number 11/926449 was filed with the patent office on 2009-04-30 for lean premixed, radial inflow, multi-annular staged nozzle, can-annular, dual-fuel combustor.
This patent application is currently assigned to General Electric Company. Invention is credited to Gregory A. Boardman, Thomas E. Johnson, Johnie F. McConnaughhay, Anuradha Sanyal.
Application Number | 20090111063 11/926449 |
Document ID | / |
Family ID | 40490416 |
Filed Date | 2009-04-30 |
United States Patent
Application |
20090111063 |
Kind Code |
A1 |
Boardman; Gregory A. ; et
al. |
April 30, 2009 |
LEAN PREMIXED, RADIAL INFLOW, MULTI-ANNULAR STAGED NOZZLE,
CAN-ANNULAR, DUAL-FUEL COMBUSTOR
Abstract
A lean premixed, radial inflow, multi-annular staged nozzle for
creating three independent combustion zones within a can-annular,
dual-fuel gas turbine combustor is provided. The nozzle includes a
pilot zone fueled by a gas pilot nozzle and center cartridge; a
flame holder zone fueled by an inner main gas fuel; a main flame
zone fueled by an outer main gas fuel; a main radial swirler for
mixing a portion of incoming air to the nozzle with the inner main
gas fuel supply and the outer main gas fuel supply; an endcover;
and means for controlling the ratio of an inner main gas fuel
supplied and an outer main gas fuel supplied.
Inventors: |
Boardman; Gregory A.;
(Greer, SC) ; Johnson; Thomas E.; (Greer, SC)
; McConnaughhay; Johnie F.; (Greenville, SC) ;
Sanyal; Anuradha; (Bangalore, IN) |
Correspondence
Address: |
GE ENERGY GENERAL ELECTRIC;C/O ERNEST G. CUSICK
ONE RIVER ROAD, BLD. 43, ROOM 225
SCHENECTADY
NY
12345
US
|
Assignee: |
General Electric Company
|
Family ID: |
40490416 |
Appl. No.: |
11/926449 |
Filed: |
October 29, 2007 |
Current U.S.
Class: |
431/8 ;
431/352 |
Current CPC
Class: |
F23R 3/36 20130101; F23R
3/343 20130101; Y02T 50/60 20130101; F23R 3/286 20130101; F23R 3/14
20130101; Y02T 50/675 20130101 |
Class at
Publication: |
431/8 ;
431/352 |
International
Class: |
F23C 5/08 20060101
F23C005/08 |
Claims
1. A lean premixed, radial inflow, multi-annular staged nozzle for
creating three independent combustion zones within a can-annular,
dual fuel gas turbine combustor, the nozzle comprising: a pilot
zone fueled by a center cartridge during liquid operation and a
center gas pilot nozzle during gas operation; a central flame
holder zone fueled by an inner main gas fuel supply; a main flame
zone fueled by an outer main gas fuel supply; a main radial swirler
for mixing a portion of incoming air to the nozzle with the inner
main gas fuel supply and the outer main gas fuel supply; means for
controlling the ratio of an inner main gas fuel supply and an outer
main gas fuel supply; and an endcover.
2. The nozzle according to claim 1, the main radial swirler
comprising: a backplate mechanically fastened in axial alignment to
the endcover; a plurality of swirl vanes spaced approximately
equidistantly in a circular array around a central axis of the
nozzle; a premixing space in a circumferential space between
individual swirl vanes; a central hub including a central cavity
within; and an annular swirl volume between the plurality of swirl
vanes and the central hub.
3. The large single radial nozzle according to claim 2, the
backplate comprising: a cylindrical plate; a central hub projecting
axially downstream from a downstream surface of the backplate,
wherein the central hub includes a smooth conical surface,
truncated at a downstream end; a cavity connecting a outer main gas
fuel supply from the endcover to the plurality of swirl vanes; a
plurality of injector nozzles mounted on the downstream surface of
the backplate; a cavity connecting a inner main gas fuel supply
from the endcover to the plurality of injector nozzles; liquid fuel
atomizers mounted on the downstream surface of the backplate; a
cavity connecting a liquid fuel supply from the endcover to the
plurality of liquid fuel atomizers; and a central cavity along the
central axis of the nozzle.
4. The nozzle according to claim 3, the cavity connecting the outer
main gas fuel from the endcover to the plurality of swirl vanes
further comprising: an orifice with the cavity for regulating outer
main gas fuel supply to each swirl vane.
5. The nozzle according to claim 3, wherein the cavity connecting
the liquid fuel supply includes a thermal insulating outer
liner.
6. The nozzle according to claim 3, the backplate further
comprising: a plurality of radial-oriented cavities connecting an
outer circumferential surface of the backplate with the central
cavity for supplying air to the center cartridge and the gas pilot
nozzle.
7. The nozzle according to claim 2, the swirl vane comprising: an
airfoil projecting axially downstream from a downstream surface of
the backplate towards a combustion end of the nozzle, wherein a
centerline of the airfoil making a predetermined angle with a
radius from the central axis of the nozzle, thereby defines the
circumferential premixing space for air flow from outside the main
swirler to an annular swirl volume between the main swirl vanes and
the central hub; an internal cavity within each airfoil for the
outer main gas fuel supply within the backplate; and a plurality of
gas fuel injector holes for distributing outer main gas fuel supply
to the premixing space from the internal cavity.
8. The nozzle according to claim 7, the airfoils further
comprising: a leading edge; a trailing edge; sidewalls tapering
from the leading edge to the trailing edge; the internal cavity
within each airfoil extending from a bottom surface of the airfoil
along the axial length of the airfoil, and being aligned radially
and circumferentially with the outer main gas fuel supply cavity in
the backplate; and the gas fuel injector holes being disposed
axially along the length of the internal cavity with the airfoil,
each individual hole extending from the internal cavity to an
opening on at least one sidewall of the airfoil for injecting outer
main gas fuel normal to the airflow between adjacent airfoils.
9. The nozzle according to claim 2, the premixing volume in
circumferential space between individual swirl vanes further
comprising: a radial-flaring portion of an outer burner tube
overhead forming a downstream roof.
10. The nozzle according to claim 9, wherein the radial flaring
portion of the outer burner tube is mechanically attached to the
top surface of the plurality of airfoils by mechanical means.
11. The nozzle according to claim 1, the endcover comprising: a
cylindrical plate, including an outer radial mounting surface for
mechanical attachment to the combustor and an inner radial mounting
surface for attachment to the backplate; a cavity connecting a
outer main gas fuel supply the backplate; a cavity connecting a
inner main gas fuel supply to the backplate; a plurality of
cavities connecting a liquid fuel supply from the endcover to the
backplate; and a central cavity, including a mounting flange, for
accepting and mounting a center gas pilot nozzle.
12. The nozzle according to claim 1, the central flameholder zone
comprising: a center hub; an inner burner tube wall; a perforated
cup including a plurality of holes about a central axis with
fillets between the holes and the perforated cup, a shroud around
an upper open-end of the perforated cup, and a circular opening in
the bottom of the cup, wherein the shroud mates with the inner
burner tube wall and the circular opening in the bottom of the cup
is open to the center gas pilot.
13. The nozzle according to claim 1, the central flameholder zone
comprising: a center hub; an inner burner tube wall; a deswirler
for converting a circumferential flow of a fuel-air mixture in the
annular swirl volume of the main radial swirler and redirecting the
airflow in an axial downstream direction; and a v-gutter flame pack
holder.
14. The nozzle according to claim 13, the deswirler comprising: a
plurality of segmented radial compartments, each compartment formed
as an annular segment bounded on an outer radius by an inner burner
tube wall and on an inner radius by an outside wall of the center
hub, wherein adjacent compartments are separated by vanes of
circumferentially-sloping radial walls wherein the slope
progressively increases from an upstream entrance to the
compartment to the downstream exit from the compartment, for
deswirling a portion of the fuel-air mixture in the annular swirl
volume.
15. The nozzle according to claim 13, the v-gutter flame holder
pack comprising: a plurality of radial oriented arms spaced
approximately equidistant circumferentially around the inner burner
tube, the arms being attached and extending from a downstream end
of the center hub to a downstream axial end of the inner burner
tube, wherein the attachment at the inner burner wall is located
downstream from the attachment at the hub extension thereby forming
a predetermined radial-axial angle for the radial oriented arms,
and a convex shaped depression in the radial-oriented arms, a
vertex of the convex shaped depression pointing upstream.
16. The nozzle according to claim 13, the center hub comprising: a
cylindrical tube with a central cavity, an irregular shape of an
inner surface of cylindrical tube to accommodate the central gas
pilot nozzle, and adapted at an upstream end to mate with the
central hub of the main radial swirler, further including a hub
extension on a downstream axial end of the cylindrical tube, the
hub extension being interrupted by v-gutters at equidistant
intervals around the hub.
17. The nozzle according to claim 3, the central gas pilot nozzle
comprising: a generally cylindrical-shaped body with a center
cavity and a radially-expanded bolting flange at an upstream end;
wherein the nozzle body is adapted to fit within a central cavity
of the endcover, the backplate and a center hub of the flameholder
zone; a plurality of radial air feed holes aligned axially on the
body to accept airflow from the central cavity of the backplate;
and a central cartridge including a liquid fuel pilot disposed
within the center cavity; an igniter disposed within the center
cavity; an end tip, providing axial and radial support on the
downstream end for the liquid fuel pilot and the igniter; a liquid
pilot fuel supply; a gas pilot fuel supply; and a power supply
connection for the igniter.
18. The nozzle according to claim 17, the central gas pilot nozzle
further comprising: an annulus formed between the central cavity of
the backplate and the center hub of the flameholder for passing a
portion of air supplied to the central cavity of the backplate to
the gas pilot nozzle and the central cartridge; an annulus,
positioned inner radially from an inside wall of the central
cartridge, for supplying pilot gas fuel to the gas pilot from a gas
pilot supply at the downstream end of the central cartridge; a
plurality of pilot swirl vanes projecting radially outward between
the outer surface of the central cartridge and the inner surface of
the center hub of the flameholder for mixing air in the annulus
with pilot gas fuel, the swirl vanes positioned equidistant around
the outer surface; a plurality of radially extending holes through
the central cartridge wall, individual holes, radially-centered and
axially-positioned at the upstream entrance between adjacent pilot
swirl vanes; and an annulus located downstream of the swirl vanes
for supplying a swirled pilot gas-air mixture to the pilot
zone.
19. The nozzle according to claim 17, the central cartridge
comprising: an axial body for the liquid fuel pilot; an axial
channel through the center of the liquid fuel pilot for liquid
fuel; an insulating annular air gap surrounding the axial body; an
air assist supply for the liquid fuel pilot; an air assist annulus
surrounding the axial body; a truncated conical liquid fuel spin
chamber; a truncated conical heat shield surrounding the liquid
fuel spin chamber; a truncated conical annulus surrounding the heat
shield and a tip impingement shield with air passage holes for
center cartridge air.
20. The nozzle according to claim 19, the central cartridge further
comprising: a center cartridge air chamber running the axial length
of the center cavity of the central cartridge and bounded on the
downstream end by the tip impingement shield; the tip impingement
shield including a plurality of axial holes; an air chamber
positioned axially between the tip impingement shield and the end
tip; a channel to the igniter from the air chamber; and a channel
to the truncated conical annulus surrounding the heat shield on the
liquid fuel pilot.
21. The nozzle according to claim 1, the main zone further
comprising: a cylindrical inner burner tube wall centered on a
central axis of the nozzle; a cylindrical outer burner tube wall
centered on the central axis of the nozzle, the outer burner tube
wall projecting axially downstream from the downstream surface of
the main radial swirler and being of larger diameter being than the
inner burner tube wall; and a flared portion of the outer burner
tube at its upstream end extending outward radially as an annular
circumferential surface, the circumferential surface forming a roof
over the plurality of main swirl vanes and channeling the fuel and
air into the annular mixing zone.
22. The nozzle according to claim 1, the means for controlling the
ratio of an inner main gas fuel supplied and an outer main gas fuel
supplied comprising: at least one of means for throttling the fuel
supplies and means for changing the pressure of the fuel
supplies.
23. A can-annular dual-fuel combustor for a gas turbine engine,
comprising: a lean premixed, radial inflow, multi-annular staged
nozzle, including an inner burner tube, an outer burner tube and a
main radial swirler mounted on an endcover to a combustor casing; a
main combustion zone downstream from the outer burner tube of the
nozzle; a source of compressed air from a compressor; an air inlet
plenum radially surrounding the single large radial nozzle and
bounded radially by an outer wall of the combustor; a diffuser for
the compressed air, the diffuser receiving the compressed air in a
reverse flow path from the compressor and discharging the
compressed air at a restored pressure to the inlet plenum; and a
fairing mounted atop the main radial swirler and surrounding a
portion of the outer burner tube for smoothing air flow from the
diffuser to the air inlet plenum.
24. The can-annular dual-fuel combustor according to claim 23, the
diffuser further comprising: an inner wall, the inner wall being
coincident with the backside for a dome on the main combustion
zone, thereby providing backside cooling to the dome from the
compressed air passing through the diffuser.
25. The can-annular dual-fuel combustor according to claim 23, the
outer burner tube further comprising: a reinforced cylindrical
segment in proximity to the downstream end of the tube, including a
cylindrical ledge wherein the cylindrical ledge provides axial and
radial support for the dome.
26. The can-annular dual-fuel combustor according to claim 23,
wherein the combustor further comprising: a plurality of supply and
control connections to an endcover including: an outer main gas
fuel supply; an inner main gas fuel supply; a gas pilot supply a
liquid fuel supply; a liquid pilot fuel supply; an air assist
supply; igniter controls; and flame detector controls.
27. A method for utilizing a lean premixed, radial inflow,
multi-annular staged nozzle with independent combustion zones
including a pilot zone, a flame holder zone and a main zone, within
a can-annular dual-fuel gas turbine combustor for providing stable
combustion with low Nitrogen Oxide (NOx) emissions, the method
comprising providing a large supply of air to the nozzle;
intra-nozzle staging; breaking up heat release into a multiplicity
of discrete zones in space; distributing the heat release in time;
and ventilating a downstream central recirculation zone.
28. The method for utilizing a lean premixed, radial inflow,
multi-annular staged nozzle within a gas turbine combustor for
providing stable combustion with low NOx emissions according to
claim 27, the step of providing a large supply of air to the nozzle
comprising: providing a channeled reverse flow path for air to the
nozzle from an air compressor; recovering static air pressure by
diffusing air flow through a diffuser tube positioned between the
reverse flow path for air and an inlet plenum for the nozzle;
smoothing air flow entering the inlet plenum for the nozzle with a
fairing shaped around the outer flame holder for the nozzle; and
flowing air through large flow passages of a single large radial
nozzle.
29. The method for utilizing a lean premixed, radial inflow,
multi-annular staged nozzle within a gas turbine combustor for
providing stable combustion with low NOx emissions according to
claim 27, the step of intra-nozzle staging comprising: providing a
outer main fuel supply; providing a inner main fuel supply; mixing
the outer main fuel supply with air from the inlet plenum at an
outboard injection point between adjacent swirl vanes of the main
radial swirler, the injection point being outboard radially of the
swirl vanes; mixing the inner main fuel supply with air from the
inlet plenum at an inboard injection points radially inboard of the
swirl vanes; controlling the ratio of fuel injected at the outboard
injection point and the inboard injection point; and biasing a
fuel-air ratio to be richer near the central hub of the main radial
swirler; and burning fuel in the flame holder zone at a higher
equivalence ratio relative to burning in the main zone.
30. The method for utilizing a lean premixed, radial inflow,
multi-annular staged nozzle within a gas turbine combustor for
providing stable combustion with low NOx emissions according to
claim 27, the step of breaking up the heat release into a
multiplicity of discrete zones in space comprising: providing three
spatially-separated burn zones including the pilot zone, the
flame-holder zone and the main zone; and employing a physical
separation of a plurality sloping radial arms of the v-gutter pack
to create a multiplicity of discrete reaction zones in space, each
reacting at spatial scales that are much smaller than that of the
overall combustion chamber effectively limiting the amount of
energy release that can constructively couple at a particular
resonant frequency within the combustion chamber.
31. The method for utilizing a lean premixed, radial inflow,
multi-annular staged nozzle within a gas turbine combustor for
providing stable combustion with low NOx emissions according to
claim 27, the step of distributing the heat release in time
comprising: establishing a discrete fuel transport time for each
point along a length of a v-gutter, effectively limiting the amount
of energy release that can constructively couple at a particular
resonant frequency within the combustion chamber.
32. The method for utilizing a lean premixed, radial inflow,
multi-annular staged nozzle within a gas turbine combustor for
providing stable combustion with low NOx emissions according to
claim 27, the step of ventilating a downstream central
recirculation zone comprising: deswirling the a fuel-air mixture,
from the main radial swirler, in the central flame-holder;
injecting non-swirling axial momentum into the central
recirculation zone; and limiting the average time that
combustion-product molecules spend at flame temperatures in the
combustion zone.
Description
BACKGROUND OF THE INVENTION
[0001] The invention relates generally to gas turbine combustors
and more specifically to a lean premixed, radial inflow,
multi-annular staged nozzle for a can-annular dual-fuel combustor
that dramatically reduces or eliminates combustion dynamics.
[0002] FIG. 1 illustrates a prior art combustor for a heavy-duty
industrial gas turbine 10, which includes a compressor 12
(partially shown), a plurality of combustors 14 (one shown for
convenience and clarity), and a turbine 16 (represented by a single
blade). Although not specifically shown, the turbine 16 is
drivingly connected to the compressor 12 along a common axis. The
compressor 12 pressurizes inlet air, which is then reverse flowed
to the combustor 14 where it is used to cool the combustor 14 and
to provide air to the combustion process. Although only one
combustor 14 is shown, the gas turbine 10 includes a plurality of
combustors 14 located about the periphery thereof. A transition
duct 18 connects the outlet end of each combustor 14 with the inlet
end of the turbine 16 to deliver the hot products of combustion to
the turbine 16.
[0003] Each combustor 14 includes a substantially cylindrical
combustion casing 24 which is secured at an open forward end to a
turbine casing 26 by means of bolts 28. The rearward end of the
combustion casing 24 is closed by an end cover assembly 30 which
may include conventional supply tubes, manifolds and associated
valves, etc. for feeding gas, liquid fuel and air (and water if
desired) to the combustor 14. The end cover assembly 30 receives a
plurality (for example, five) of fuel nozzle assemblies 32 (only
one shown for purposes of convenience and clarity) arranged in a
circular array about a longitudinal axis of the combustor 14. Each
fuel nozzle assembly 32 is a substantially cylindrical body having
a rearward supply section 52 having inlets for receiving gas fuel,
liquid fuel and air (and water if desired) and a forward delivery
section 54.
[0004] Within the combustion casing 24, there is mounted, in
substantially concentric relation thereto, a substantially
cylindrical flow sleeve 34 which connects at its forward end to the
outer wall 36 of the transition duct 18. The flow sleeve 34 is
connected at its rearward end by means of a radial flange 35 to the
combustion casing 24 at a butt joint 37 where fore and aft sections
of the combustor casing 24 are joined.
[0005] Within the flow sleeve 34, there is a concentrically
arranged combustion liner 38, which is connected at its forward end
with the inner wall 40 of the transition duct 18. The rearward end
of the combustion liner 38 is supported by a combustion liner cap
assembly 42 which is, in turn, supported within the combustion
casing 24 by a plurality of struts 39. It will be appreciated that
the outer wall 36 of the transition duct 18, as well as that
portion of flow sleeve 34 extending forward of the location where
the combustion casing 24 is bolted to the turbine casing 26 (by
bolts 28) are formed with an array of apertures 44 over their
respective peripheral surfaces to petit air to reverse flow from
the compressor 12 through the apertures 44 into the annular space
between the flow sleeve 34 and the liner 38 toward the upstream or
rearward end of the combustor 14 (as indicated by the flow arrows
shown in FIG. 1).
[0006] The combustion liner cap assembly 42 supports a plurality of
premix tubes 46, one for each fuel nozzle assembly 32. More
specifically, each premix tube 46 is supported within the
combustion liner cap assembly 42 at its forward and rearward ends
by front and rear plates 47, 49, respectively, each provided with
openings aligned with the open-ended premix tubes 46. The premix
tubes 46 are supported so that the forward delivery sections 54 of
the respective fuel nozzle assemblies 32 are disposed
concentrically therein.
[0007] The rear plate 49 mounts a plurality of rearwardly extending
floating collars 48 (one for each premix tube 46), arranged in
substantial alignment with the openings in the rear plate 49. Each
floating collar 48 supports an annular air swirler 50 in
surrounding relation to the respective fuel nozzle assembly 32.
Radial fuel injectors 66 are provided downstream of the swirler 50
for discharging gas fuel into a premixing zone 69 located within
the premix tube 46. The arrangement is such that air flowing in the
annular space between the liner 38 and the flow sleeve 34 is forced
to again reverse direction in the rearward end of the combustor 14
(between the end cap assembly 30 and sleeve cap assembly 42) and to
flow through the swirlers 50 and premix tubes 46 before entering
the burning zone or combustion chamber 70 within the liner 38,
downstream of the premix tubes 46. Ignition is achieved in the
multiple combustors 14 by means of a spark plug 20 in conjunction
with cross fire tubes 22 (one shown) in the usual manner.
[0008] In power plant design, reducing emissions of harmful gases
such as nitrogen oxides (NOx) into the atmosphere is of prime
concern. Low NOx combustors employing lean premixed combustion with
a plurality of burners attached to a single combustion chamber,
such as described in FIG. 1, have been developed to approach this
problem. Each burner includes a flow tube with a centrally-disposed
fuel nozzle comprising a cylindrical hub, which supports fuel
injectors and an air swirler and has a flat face on its downstream
end. In addition to a premix-injection stage for low NOx operation,
each fuel nozzle can include a diffusion-injection stage for
start-up and emergency operations and a liquid fuel-injection stage
for liquid fuel operation.
[0009] Diffusion gas fuel and liquid fuel are typically injected
via orifices located on the flat end face of the fuel nozzle.
During low NOx (premix) operation, fuel is injected through the
fuel injectors and mixes with the swirling air in the flow tube.
The diffusion and liquid fuel circuits are typically purged with
air during premix operation to keep flame gases out of the
passages. The combustion flame is stabilized by bluff-body
recirculation behind the fuel nozzle and swirl breakdown, if swirl
is present. With premixed systems, strong pressure oscillations are
typically produced as a result of combustion instabilities. The
combustion instabilities are believed to be related to the shedding
of spanwise vortices from the bluff end of the fuel nozzle. These
pressure oscillations can severely limit the operation of the
device and in some cases can even cause physical damage to
combustor hardware. Furthermore, the flow of purge air through the
diffusion and liquid fuel circuits is injected directly into the
recirculation zone. This direct injection reduces the local
temperature and strength of the recirculation, producing an adverse
effect on flame stability. Accordingly, there is a need for a low
NOx combustor, which reduces pressure oscillations and avoids the
adverse effects of injecting purge air directly into the
recirculation zone.
[0010] As previously described, these contemporary heavy-duty
industrial Dry-Low NOx (DLN) can-annular gas turbine combustors
typically employ a multiplicity (or gang) of premixing nozzles
interfaced with a can combustor liner using a flat or angled
cap/dome assembly. Multiple nozzles are required for the mixing and
the staging of fuel to achieve turndown and performance throughout
the intended operability and design space. This approach, however,
creates a complicated and expensive assembly.
[0011] Also, distributing the air and fuel uniformly to the cluster
of premixing fuel nozzles at the headend is difficult and generally
results in less than ideal, uniform air flow to all the nozzles, or
a substantial amount of parasitic pressure drop/loss.
Swirl-stabilized, lean premixed (LP) combustion tends to be highly
susceptible to combustion-driven oscillations (dynamic instability)
compared to conventional, diffusion style combustion.
[0012] Historically, in the gas-turbine-engine industry, flame
temperature (or primary zone temperature) has been reduced in LP
systems to reduce NOx emissions. As acceptable NOx exhaust
emissions levels have been decreased down to single digit
parts-per-million (ppm) levels (driven primarily by new government
regulations) flame temperature has been driven very near to the
lean-blowout (LBO) limit, at least for fuels with a high methane
content. For such lean mixtures, slight, periodic variations in
local fuel-to-air mixture ratio results in relatively large,
periodic variations in local heat release and heat-release
rates--even including local, periodic flame extinction. Discrete,
oscillation frequencies (or tones) can grow in amplitude when the
heat-release fluctuations are constructively in phase with the
acoustic pressure fluctuations encountered inside the combustion
chamber.
[0013] As present LP combustors become leaner and more spatially
uniform to meet increasingly lower NOx emissions, and are
increasingly required to meet those emission targets while running
on a broadening range of fuels, the risk of encountering
unacceptably high levels of combustion dynamics goes up for a given
system.
[0014] Although, large single-nozzle DLN, low-NOx can-annular
gas-turbine combustion systems have been tried previously, most
have failed due to operability, durability, and emissions problems.
Lack of smart, tunable operating parameters and a lack of multiple
independent combustion staging zones has led the industry to
embrace modular, multi-nozzle (gang) configurations. Multi-nozzle
designs allow for the staging or skewing of fuel distribution to
subgroups of nozzles to not only facilitate lightoff and turndown,
but to provide a tunable operability parameter to skirt dynamics
(or oscillations) encountered while running in the design,
operating space.
[0015] The downside of skewing the fuel distribution in the
combustor is that hotter temperature zones are created that drive
NOx production. Thus, if too much skewing is required to squash
dynamics or instability, the breaching of regulatory NOx emissions
limits could occur, possibly putting the unit out of commission. LP
combustion dynamics in industrial gas turbines are typically abated
passively in a few ways, usually a trial and error process, which
can be expensive and uncertain. Some of the methods are listed
below: 1) shifting the fuel injection points to alter the fuel
transport time from the point of injection to the flame front, 2)
changing the fuel injection orifice sizes to alter the pressure
drop and the acoustic impendence across the holes, and 3) modifying
chamber or nozzle geometries (e.g., diameters, angles, lengths) to
affect vortex shedding, frequencies and amplitudes, or flame shape
in the chamber.
[0016] These methods attempt to force any perturbations in heat
release to be out of phase (or destructively in phase) with
pressure or acoustic perturbations in the combustion chamber.
Combustor dynamics have also been reduced or eliminated by adding
acoustic damping (e.g., Helmholtz resonators or quarter-wave tubes)
to the combustion system. In the past, the above methods have
tended to be considered and exercised ex post facto to discovering
high combustor dynamics, instead of designing for them proactively
during the initial design phase in the program.
[0017] Accordingly, there is a need to provide a simpler, scalable,
less-expensive LP combustor that is fundamentally much less likely,
in a statistical and absolute sense, to excite or drive discrete
combustion oscillations at any loading within the design/operating
space, while having an above average tolerance to fuel-mixture
quality. Assuming that if the above solution were found, and that,
consequently, the risk of ever encountering discrete dynamics in
the given design operating space were greatly reduced, then the
efficiency and probability of tuning for the minimum emissions for
a given system would be greatly increased. Essentially, dynamics
would no longer be such a significant and intractable part of the
overall combustor-design procedure.
BRIEF DESCRIPTION OF THE INVENTION
[0018] The present invention relates to an apparatus and method
creating three independent combustion zones in a gas turbine
combustor with a lean premixed, radial inflow, multi-annular,
staged nozzle, thereby providing stable combustion with low
nitrogen oxide (NOx) emissions.
[0019] Briefly in accordance with one aspect of the present
invention, a lean premixed, radial inflow, multi-annular, staged
nozzle for creating three independent combustion zones within a
can-annular, dual-fuel gas turbine combustor is provided. The lean
premixed, radial inflow, multi-annular, staged nozzle (hereinafter
referred to as a single large radial nozzle) includes a pilot zone
fueled by a center cartridge; a flame holder zone fueled by an
inner main gas fuel; a main flame zone fueled by an outer main gas
fuel; a main radial swirler for mixing a portion of incoming air to
the nozzle with the inner main gas fuel supply and the outer main
gas fuel supply; an endcover; and means for controlling the ratio
of pilot gas fuel supplied, inner main gas fuel supplied, and an
outer main gas fuel supplied.
[0020] In accordance with another aspect of the present invention,
a can-annular, dual-fuel combustor for a gas turbine engine is
provided. The combustor includes a lean premixed, radial inflow,
multi-annular, staged nozzle (hereinafter referred to as a single
large radial nozzle), incorporating an outer burner tube and a main
radial swirler, mounted on an endcover to a combustor casing. A
main combustion zone is provided downstream from the outer burner
tube of the single large radial nozzle. A source of compressed air
from a compressor source is provided. An air inlet plenum radially
surrounds the single large radial nozzle and is bounded radially by
an outer wall of the combustor. A diffuser for the compressed air
receives the compressed air in a reverse flow path from the
compressor and discharges the compressed air at a restored pressure
to the inlet plenum. A fairing mounted atop the main radial swirler
and surrounding a portion of the outer burner tube is provided for
smoothing air flow from the diffuser to the air inlet plenum.
[0021] In accordance with a third aspect of the present invention,
a method is provided for utilizing a lean premixed, radial inflow,
multi-annular, staged nozzle (hereinafter referred to as a single
large radial nozzle) with independent combustion zones, wherein the
single large radial nozzle includes a pilot zone, a flame holder
zone and a main zone, within a gas turbine combustor for providing
stable combustion with low Nitrogen Oxide (NOx) emissions. The
method includes providing a large supply of air to the nozzle;
intra-nozzle staging; breaking up of the heat release into a
multiplicity of discrete zones in space; distributing the heat
release in time; and ventilating a downstream central recirculation
zone.
BRIEF DESCRIPTION OF THE DRAWING
[0022] These and other features, aspects, and advantages of the
present invention will become better understood when the following
detailed description is read with reference to the accompanying
drawings in which like characters represent like parts throughout
the drawings, wherein:
[0023] FIG. 1 illustrates a prior art combustor with multiple
nozzles;
[0024] FIG. 2 illustrates an embodiment of an inventive combustor
including an inventive large single radial nozzle combustor;
[0025] FIG. 3A illustrates an isometric cutaway showing an internal
structure for an embodiment of the inventive large single radial
nozzle structure;
[0026] FIG. 3B illustrates an axial cross-section showing an
internal structure for an embodiment of the inventive large single
radial nozzle structure;
[0027] FIG. 4 illustrates an supply-end view of the endcover plate
for an embodiment of the inventive large single radial nozzle:
[0028] FIG. 5 illustrates fuel galleries and fuel penetrations in
the endcover and backplate of an embodiment of the inventive large
single radial nozzle;
[0029] FIG. 6A illustrates an isometric view for a main radial
swirler of an embodiment of the inventive large single radial
nozzle;
[0030] FIG. 6B illustrates details of main swirl vanes on the main
radial swirler of an embodiment of the inventive large single
radial nozzle;
[0031] FIG. 6C illustrates an elevation view of the main radial
swirler for an embodiment of the inventive large single radial
nozzle;
[0032] FIG. 6D illustrates a section through the main swirl vanes
and central hub of the main radial swirler for an embodiment of the
inventive large single radial nozzle
[0033] FIG. 6E illustrates a section through the backplate and
central hub of the main radial swirler of an embodiment of the
inventive large single radial nozzle;
[0034] FIG. 7 illustrates a cross-section of the head end of the
inventive combustor depicting air flow and independent combustion
zones of an embodiment of the inventive large single radial
nozzle;
[0035] FIG. 8 illustrates the central flame holder, gas pilot
annulus, and center cartridge of an embodiment of the inventive
large single radial nozzle;
[0036] FIGS. 9A and 9B illustrate the nozzle body of the gas pilot
nozzle for an embodiment of the inventive large single radial
nozzle;
[0037] FIG. 10 illustrates an axial section view of a dual-fuel
center cartridge of an embodiment of the inventive large single
radial nozzle; and
[0038] FIG. 11 illustrates an alternate embodiment for the
flameholder cup for the inventive large single radial nozzle.
DETAILED DESCRIPTION OF THE INVENTION
[0039] The following embodiments of the present invention have many
advantages, including several innovative and unique features: (1)
allowing for multiple (e.g., six) premixing nozzles (per can) and a
combustor-chamber cap to be replaced with just one large radial
nozzle and a liner modification, thereby achieving a significant
part-count reduction, a cost savings, and a dramatic simplification
of the combustor's head-end; (2) using a dome-diffuser design to
backside, convectively cool the liner's dome, while, simultaneously
recovering static pressure prior to premixing the fuel and air in
the large radial nozzle, thereby causing less parasitic pressure
loss and malting more air available for premixing; (3) providing
the capacity to rapidly (e.g. <3 msec) and thoroughly vaporize
and mix large quantities of fuel (.about.2 lbm/sec) and air
(.about.60 lbm/sec) at a relatively low pressure drop (e.g.,
<4%); and (4) using either gas fuel or liquid fuel, it is more
robust dynamically and less prone to combustion-driven oscillations
than contemporary lean-premixed (LP) gas-turbine combustion
systems, by strategically distributing or smearing out both the
heat release (in time and space) and fuel transport time in the
chamber, while still delivering the necessary system turn-down
performance and low exhaust emissions.
[0040] The effect of the design is multifaceted: (1) the
replacement of five or more nozzles per can with one for a dramatic
cost and parts-count reduction; (2) the dramatic reduction or even
complete elimination of combustion dynamics/oscillations at
discrete frequencies within industrial gas turbine combustion
chambers, while maintaining required emissions levels; (3) gas and
liquid fuels flexibility, which is tied to the success of
combustion dynamics improvement; (4) DLN with liquid fuels like No.
2 diesel oil, while eliminating the need for water injection and
high-pressure atomizer air; and (5) low single-digit (ppmv) noxious
emissions.
[0041] To go from a multi-nozzle arrangement to a large single
nozzle, successfully, requires intra-nozzle staging. A zone of
angled v-gutter flame holders, used in this design, provides a
region for fuel staging within the main premixing nozzle. For
example, biasing the fuel-air ratio to be richer near the hub (or
centerbody) in the premixer, can allow a central flame-holder zone
to burn at a higher equivalence ratio relative to the bypassing
flow, which may be advantageous (or even necessary) for ignition,
machine acceleration, low-load operation, or handling sudden load
transfers. Biasing the fuel-air ratio, in conjunction with other
staging features (like a premixed pilot), will allow for a single
large radial nozzle to replace multiple nozzles (e.g., six), per
can, in gas turbine combustion systems--which would amount to a
significant part-count reduction and cost savings for the
combustion system and the engine as a whole. Combustion dynamics
reduction would be achieved while maintaining or even reducing
noxious exhaust emissions (e.g., unburnt hydrocarbons (UHC), NOx,
and carbon monoxide (CO)), relative to the firing temperatures
encountered in the design space.
[0042] The lean premixed, radial inflow, multi-annular staged
nozzle (hereinafter referred to as a single large radial nozzle),
by design, is less likely to excite combustion-driven, discrete
oscillation frequencies when the set fraction (e.g., about 33% of
the nozzle's reactants) of well-mixed reactants is redirected to
burn as an angled array of discrete, v-gutter zones upstream of the
main chamber. The array of axial jets passing through the conical
v-gutter-pack structure abates discrete dynamics and improves
emissions in a few ways.
[0043] First, the array breaks up the heat release into a
multiplicity of discrete reaction zones in space, each reacting at
spatial scales that are much smaller than that of the overall
combustion chamber. This effectively limits the amount of energy
release that can constructively couple at a particular acoustic,
resonant frequency within the chamber.
[0044] Second, the angled v-gutters create a multiplicity of fuel
transport times, which distributes (or smears out) the heat release
in time. That is, each point along the v-gutter length has its own
associated transport time: the time between the point(s) of fuel
injection and the point(s) of burning. This, too, effectively
limits the amount of heat-release energy that can constructively
couple at a particular chamber acoustical, resonant frequency.
[0045] Third, the function of the de-swirler pack is the
"ventilation" of the downstream central recirculation zone (CRZ),
resulting from vortex breakdown. From the central cone, the
expanding array of jets injects non-swirling, axial momentum
directly into the CRZ, which reduces the size and the bulk
residence time of the CRZ. This, in turn, reduces nitrogen oxide
(NOx) production by reducing the average time that the
combustion-product molecules spend at the primary-zone (flame)
temperature inside the combustor. The concept of "time at
temperature" for NOx production becomes increasingly significant at
flame temperatures above 2900 F, where the Thermal NOx (or
Zeldovich) mechanism begins to accelerate and its contribution to
overall system NOx levels begins to significantly increase.
[0046] The nozzle further provides an anti-coke design for
operation with diesel liquid fuel that requires no water and no
atomizer air. Aspects of the design prevent fuel gallery coking
through an insulated fuel gallery wall for high reliability. Liquid
fuel is rapidly atomized and thoroughly dispersed into the premixer
airflow, keeping it off of hot premixer surfaces to vaporize and
mix quickly with the air. The liquid injection scheme does not
adversely impact gas operation. Eliminating the need for water
injection and high-pressure atomizer air further provides a cost
saving.
[0047] Complete fuel-air mixing is rapid (approximately 2 msec),
thorough (greater than 97%), and requires a low premixer
differential pressure (.about.2%), thereby reducing the required
premixer residence time to create a shorter, more compact design
and to stay below the auto-ignition time of diesel at "advanced"
gas turbine conditions.
[0048] Several further aspects and advantages of the invention will
become clear in the description. Turndown capability is enhanced
via fuel staging (3 pseudo-independent combustion zones). Using a
backside-cooled dome eliminates the need for liner cooling air in
the flame zone.
[0049] Also, axi-symmetric, radial combustion staging does not
subject the combustor liner to asymmetric loading, thereby
providing improved combustion liner durability.
[0050] Further, improved internal premixer flame-holding
resistance/margin: flow is accelerated throughout premixer nozzle;
bulk velocity is kept above about 300 ft/sec.
[0051] A v-gutter lean angle (radial-axial plane) and the
de-swirler vane profile were chosen as the two parameters to
optimize. The v-gutter lean angle was varied between 30- and
60-deg. to maximize the lean angle, while still generating a well
defined, continuous v-gutter wake to support an independent
combustion zone. For non-reacting CFD, the 40-deg. configuration
was the largest angle that still produced a continuous v-gutter
wake with other nozzle features being held constant. The de-swirler
vane profile was successfully adjusted/optimized by aligning the
inlet vane angle with the incoming swirling flow and using the
cascade geometry to accelerate the flow through the pack; thus,
preventing any flow separation in the pack.
[0052] FIG. 2 illustrates an embodiment of the inventive large
single radial nozzle employed in an inventive combustor. The large
single radial nozzle combustor 100 includes a substantially
cylindrical combustor casing 105, which may be secured at an open
forward end to a transition piece (not shown) for connection with a
turbine by plugging a combustion liner into the transition piece.
The transition piece may then be secured at an open forward end to
a turbine casing by bolts in the usual manner. The rearward end of
the combustor casing is closed by an endcover assembly 130
adaptable to conventional packages for supply tubes, manifolds,
valves and fittings for gas fuel, liquid fuel, air, and power (not
shown). The endcover assembly 130 is part of and secures a large
radial nozzle assembly 120 to the combustor casing 105.
[0053] Within the combustion casing 105, there is mounted in
substantially concentric relation thereto a flow sleeve 106. Within
the flow sleeve 106, there is concentrically arranged combustion
liner 110, connected at its forward end 112 to the inner wall of
the transition liner (not shown), into which it plugs. The rearward
end of combustion liner 110 forms a truncated conical dome 111 on a
main combustion chamber 114, the truncated conical dome 111 being
open at its center to fuel and combustion products flow from the
large radial nozzle 120 and further mating with the outer burner
tube 113 of the large radial nozzle 120.
[0054] Air for the combustion process may be drawn from the air
compressor into the transition piece (as previously described with
respect to FIG. 1) and thence through the annular space 115 between
the flow sleeve and the outside wall of the combustion liner. At
the aft end of the annular space, a concentric mounted diffuser 116
expands the air into the inlet plenum 117 for the large radial
nozzle 120. The dome-111 serves as an inner concentric wall of the
diffuser 116, thereby permitting backside cooling of the dome 111
by the air flowing through the diffuser 116. Simultaneously, the
diffuser 116 recovers static pressure for the air prior to
premixing the fuel and air in the large radial nozzle 120,
resulting in less parasitic pressure loss and more air being
available for premixing. A fairing 118 around the center of the
large radial nozzle 120 smoothes air entry into the inlet plenum
117 further reducing parasitic pressure loss.
[0055] The large radial nozzle 120 further includes a main radial
swirler 140, a gas pilot nozzle 150, a central flame holder with a
v-gutter flame holder 160, and an outer flame holder 170. The
central flame holder 160 and the outer flame holder 170 open on
their forward end into the main combustion chamber 114.
[0056] The endcover 130 may be a generally cylindrical-shaped
flange designed to mate with a combustor casing 105 and support the
radial nozzle assembly 120 within the combustor 100. The aft
surface 135 of the endcover 130 provides penetrations for dual-fuel
(gas and liquid fuel) as well as for the gas pilot nozzle 150. A
outer main gas supply 190, a inner main gas supply 190 and one of a
plurality of liquid gas connections 195 are shown in FIG. 3A. The
arrangement of the penetrations, for fuel and air, permit
connection to existing combustor configurations of fuel, air and
power lines (not shown).
[0057] FIG. 3A illustrates an isometric cutaway showing an internal
structure for an embodiment of the inventive large single radial
nozzle. FIG. 3B illustrates an axial cross-section showing an
internal structure for an embodiment of the inventive large single
radial nozzle. The nozzle is axi-symmetric along a central axis
200.
[0058] The endcover assembly 130 includes an endcover plate 205
with an aft section 201, a forward section 202 and a central cavity
203. A main radial swirler 140 includes a backplate 240, a
plurality of swirl vanes 250 and a central hub 260 with a central
cavity 265 within. The backplate 240 is bolted at mounting surface
241 to endcover plate 205.
[0059] A central flame holder 160 is mounted atop the central hub
260. A center hub 285 of the central flame holder 160 mates with
the central hub 260 of the main radial swirler 140 to support the
central flame holder 160 radially and axially. Radial vanes 360
support inner burner tube 300 from center hub 285. A plurality of
v-gutters 290 extend between the inner burner tube 300 and the
center hub 285. A central cavity 278 is formed within center hub
285. Atop the swirl vanes 250 of the main radial swirler 140, an
outer flame holder 170 with cylindrical outer burner tube 175 is
mounted with a base section 180 that flares outward radially and
attaches with bolts 183 to the top of the swirl vanes 250. The
downstream end 178 of outer burner tube 175 also flares outward and
is reinforced to provide support for the conical dome 111 (FIG. 2)
of the combustor. Support ledge 179 mates with combustor conical
dome 111. A fuel-air mixture from the main radial swirler 140
passes to a flameholder zone through 402 and to the main zone
through 405 (FIG. 7). The central cavities 203, 265 and 278 allow
for insertion of a gas pilot nozzle 150 including a gas pilot and a
center cartridge including, a liquid pilot and an igniter.
[0060] FIG. 4 illustrates an aft-end view of the endcover plate of
an embodiment of the inventive large single radial nozzle with
respect to a combustor in which it may be installed. The endcover
plate 205 includes the integral cylindrical shaped aft section 201
and a cylindrical shaped forward section 202 (FIG. 3A) of a smaller
diameter, both layers being centered on the central axis 200 of the
nozzle. The aft section 201 is sized radially to mate with a rear
seating surface of the combustor (not shown) and incorporates a
plurality of bolt holes 206 formed axially and in proximity to the
outer circumferential surface 207 of the outer section 201 for
attaching to the seating surface of the combustor. The aft section
201 may also include a plurality of pilot holes 208 (similarly
directed) for positioning the aft section 201 with respect to the
seating surface of the combustor, in preparation for bolting. The
forward seating surface 205 of the forward section 202 may also
include a plurality of bolt holes 209 in a circular configuration
concentric with the central axis 200 of the nozzle and adapted to
receive bolts from the mounting surface 241 of the backplate 240
for the main radial swirler 250 (FIG. 3A).
[0061] Two independent gas fuel supplies may be connected to the
endcover plate 205. The aft section 201 includes an outer main gas
penetration 215 attached to an outer main gas supply inlet pipe 216
with outer main gas inlet flange 217 for connection to the outer
main gas fuel supply (not shown). The aft section 201 also includes
an inner main gas penetration 220 with a fitting 219 for connection
to an inner main gas supply (not shown). The endcover plate 205 may
also include a plurality of liquid fuel supply penetrations 243
located concentric to the central axis 200 of the nozzle.
[0062] FIG. 4 also illustrates the aft end of gas pilot nozzle 150.
The central cavity 203 is defined within the endcover plate 205
extending radially from the central axis 200 and extending through
the aft section 201 and the forward section 202 for insertion of a
gas pilot nozzle. The central cavity 203 includes a gas pilot
nozzle seating surface 210 (FIG. 3B) with threaded connections for
mating with a gas pilot nozzle rear flange 212 to mount the gas
pilot nozzle 150
[0063] FIG. 5 illustrates fuel galleries in the endcover and
backplate of an embodiment of the inventive nozzle. The outer main
gas penetration 215 (FIG. 4) is connected to a outer main gas
gallery 310 in the endcover plate 205. The outer main gas gallery
310 defines an annular chamber concentric with the central axis 200
of the nozzle. The inner wall 311 and outer wall 312 of the outer
main gas gallery 310 may be radially located with respect to the
central axis 200 of the nozzle, such that the open upper end 315 of
the outer main gas gallery 310 communicates with a plurality of
corresponding outer main gas channels 665 (FIG. 3B) within the main
swirler backplate 240.
[0064] The inner main gas penetration 220 is connected to a inner
main gas gallery 330 in the endcover plate 205. The inner main gas
gallery 330 defines an annular chamber concentric with the central
axis 200 of the nozzle. The inner wall 317 and outer wall 318 of
the inner main gas gallery 330 may be concentric to the central
axis 200 of the nozzle. The inner main gas gallery 330) is located
radially between the outer main gas channel 310 and the central
cavity 203. The inner wall 317 and outer wall 318 of the inner main
gas gallery are radially located, such that the open upper end 319
of the inner main gas gallery 330 communicates with corresponding
inner main gas channels 680 (FIG. 7) within the main swirler
backplate 240 to supply inner main gas to inner main gas injection
points 695 on the base surface 242 between the swirl vanes 250.
[0065] Liquid fuel supply penetrations 243 extend axially through
the aft section 201 of the endcover 205, communicating with a main
liquid fuel gallery 244. The main liquid fuel gallery 244 defines
an annular chamber concentric with the central axis 200 of the
nozzle and sealed except for liquid fuel supply penetrations 243
and liquid fuel delivery penetrations 246. The main liquid fuel
gallery 244 is located radially to align with the liquid fuel
supply penetrations aft 243 and the liquid fuel delivery
penetrations 246 forward on the endcover plate 205. The liquid fuel
delivery penetrations 246 extend through the forward section 202 of
the endcover 205 to mate with corresponding liquid fuel delivery
penetrations 247 in the main radial swirler backplate 240 leading
to atomizers 248 for the liquid fuel in the main swirler backplate
240. The walls of the main liquid fuel gallery 244 and the liquid
fuel supply penetration and liquid fuel delivery penetration 246 in
the endcover 205 and fuel delivery penetrations 247 in the
backplate 240 may be provided with an insulated lining 249 to keep
wall temperatures below 290 degrees F. where coking of diesel
liquid fuel begins. Fittings 218 are provided external to the
liquid fuel supply penetrations for connection to the liquid fuel
supply.
[0066] Because the endcover plate 205 and the main swirler
backplate 240 mate in a metal to metal seating surface 241,
provision is made to isolate potential leakage from the fuel
cavities along the seating surfaces 204, 241. Three annular
recesses (FIG. 5) concentric with the nozzle central axis 200 may
be provided on the upper seating surface 204 of the endcover plate
205. The first recess 381 is provided outboard of the outer main
gas gallery 310. The second recess 382 is provided between the
outer main gas gallery 310 and the inner main gas gallery 330. The
third recess 383 is provided inboard of the inner main gas gallery
330. The recesses may be provided with C-rings or other suitable
gasketing material to prevent flow along the seating surface.
[0067] FIGS. 6A-6E illustrate views of a main radial swirler for an
embodiment of the inventive nozzle. FIG. 6A illustrates an
isometric view of a main radial swirler of an embodiment of the
inventive nozzle. FIG. 6B illustrates details of main swirl vanes
on the main radial swirler of an embodiment of the inventive
nozzle. FIG. 6C illustrates an elevation view of the main radial
swirler for an embodiment of the inventive nozzle. FIG. 6D
illustrates a section through the main swirl vanes and central hub
of the main radial swirler for an embodiment of the inventive
nozzle. FIG. 6E illustrates a section through the backplate and
central hub of the main radial swirler of an embodiment of the
inventive nozzle.
[0068] The main radial swirler 140 includes a backplate 240 with an
integral central hub 260, a plurality of main swirl vanes 250
mounted on the backplate 240 and projecting orthogonally to the
backplate 240 (downstream toward the combustion zones), a central
cavity 265 to accommodate the gas pilot nozzle 150 and a series of
internal passages within the backplate and main swirl vanes 250 to
provide for flow of fuel and air.
[0069] The backplate 240 comprises a cylindrical shaped flange
centered on the central axis 200 of the nozzle. A base surface 241
of the backplate 240 is sized radially to mate with the forward
surface 204 of the endcover plate 205. The mounting surface 242 of
the backplate incorporates a plurality of recesses 371
accommodating bolt holes 372 around the periphery the backplate
240. The bolt holes 372 extend through to the base surface 241 of
the backplate 240 and align with the bolt holes 209 on the forward
204 surface of the endcover plate 205. The mounting surface 242 of
the backplate 240 provides for mounting a plurality of main swirl
vanes 250 and housing injection points for fuel into an air flow
steam within the main radial swirler 140
[0070] The plurality of main swirl vanes 250, each including a
solid metal airfoil 610, may be mounted orthogonal to the backplate
240 and project axially toward the combustion zones. The main swirl
vanes 250 may be mounted inboard radially from the peripheral bolt
hole recesses 371 and outboard radially from the central hub 260. A
leading edge 615 of each airfoil projects generally outward
radially and a trailing edge 620 projects generally inward
radially. The axis 625 of each airfoil may form a predesignated
acute angle .alpha. (approximately 15.degree.) 630 with a radius
635 from the central axis 200 of the nozzle. While the leading edge
615 of the airfoil 610 forms a curved surface, the side surfaces
640, 641 of the airfoil 610 may form a straight-fine taper to the
common linear trailing edge 620. The bottom surface 645 and top
surface 650 of the airfoil 610 form plane surfaces. The bottom
surface 645 may be mounted to the mounting surface 242 of the
backplate 240 by welding or other suitable process.
[0071] A plurality of injection points 655 for outer main gas fuel
are provided along a radius concentric with the central axis 200 of
the nozzle, on one side surface 640 of the airfoil 610, just
inboard of the curved leading edge 615. The injection of outer main
gas fuel is provided approximately normal to the airflow 660
passing between the adjacent swirl vanes. However, injection points
655 may also be provided on both side surfaces of the airfoil and
at other locations than included in the present embodiment. The
injection points may be approximately evenly spaced axially along
the side surface 640 of the airfoil 610 to permit even distribution
of the outer main gas fuel into the airflow 660 between the
airfoils 610 in a circumferential premixing space 605. The airfoils
610 further include an internal fuel cavity 665 supplying the
injection holes 657. The fuel cavity 665 may be a generally
cylindrical-shaped hole rising from the base surface 241, axially
into the airfoil 610 and stretching in proximity to and
communicating with the injection holes 657. The fuel cavity 665
directs outer main gas fuel from the fuel cavity 310 in the
endplate 205. The injection holes 657 within each airfoil 610
extend in a radial direction with respect to the cylindrical fuel
cavity 665 to supply fuel to the injection points 655. The top
surface 650 of each airfoil 610 may further include a tapped hole
670 for securing the outer burner tube 175 to the main swirl vanes
250.
[0072] Inner main gas penetrations 680 (FIG. 7) in the backplate
240 extend axially towards the mounting surface 242 of the
backplate 240 from the inner main gas gallery 330 in the endcover
plate 205. An orifice 685 may be provided in each of the inner main
gas penetrations 680 to control gas release. Injection points 690
are provided on the mounting surface 242, approximately equidistant
from the side surfaces 640, 642 of adjacent airfoils 610 and at a
point approximately half-way along the side surface 640, 642 of the
adjacent airfoils 610. With an exemplary 24 airfoils 610, 24
injection points 690 are provided. An injection tip 695 at each
injection point 690, extending slightly above the mounting surface
242 of the backplate 240, causes the gas to be injected above the
laminar air flow on the mounting surface.
[0073] During gas operation as described above, gas fuel is
injected into the air flow of the main radial-swirler 140 from a
multiplicity of injection points 655 located axially along the
sidewalls 640 of the airfoils 610 and from injection points 695 on
the mounting surface 242 of the backplate 240. The main gas fuel is
fed from two independent feed sources as shown in FIG. 4 in order
to affect the fuel-air mixture radial profile in an annular swirl
volume (premixer annulus) 255. That is, the mixture near the
central hub 260 which eventually passes through the central flame
holder device, can be made richer or leaner compared to the mixture
near the swirl vanes 250 (which bypasses the central flame holder)
by varying the ratio of the fuel supply from the two sources.
External means may be supplied to control this ratio of an inner
main gas fuel supplied and an outer main gas fuel supplied. This
may include control throttling, pressure control or other means
known in the art, which may be exercised external to the
nozzle.
[0074] A plurality of liquid fuel injection points 245 are also
provided on the mounting surface 242 of the backplate 240 for
operation on liquid fuel. The liquid fuel injection points 245 are
positioned atop the liquid fuel delivery channels 246 in the
backplate 240. The liquid fuel channels 246 in the backplate 240
may include a thermal insulating layer 249. The liquid fuel
injection points 245 are concentric with the central 200 axis and
may be positioned to inject liquid fuel in the annular swirl volume
255 at approximately the locus of the trailing edges 620 of the
airfoils 610. In an exemplary arrangement, six liquid fuel
injection points 245 are provided circumferentially equidistant
around the mounting surface 242. Each liquid fuel injection point
245 is provided with a tip 252 that includes an atomizer 248 of a
conical shape, that screws into threads 253 for the liquid fuel
delivery channel 247. The atomizer 248 sprays liquid fuel into the
air flow in an axial direction normal to the mounting surface
242.
[0075] FIG. 7 illustrates a cross-section of the head end of the
inventive combustor depicting air flow and fuel-air flow
establishing the independent combustion zones of an embodiment of
the inventive nozzle. As previously described and referring to
FIGS. 5-7, the inventive combustor incorporating the large single
radial nozzle provides three independent combustion zones. The
pilot gas nozzle 150 creates pilot combustion zone Z1. Flame holder
combustion zone Z2 is created by the axial flow from the deswirler
280 passing over the v-gutters 290 in the central flame holder 160.
Main combustion zone Z3 is created by fuel-air mixture flowing
between the inner burner tube 300 of the central flame holder 160
and the outer burner tube 175 of the outer flame holder 170 into
the main combustion chamber 114.
[0076] Airflow from diffuser 116 flows into inlet plenum 117. The
main swirl vanes 250 establish a flow path 660 for incoming air
from the inlet plenum 117 for the combustor. About 95% of the air
entering the nozzle flows between the main swirl vanes 250. The
incoming air, having had outer main gas injected from the airfoils
610 and inner main gas injected from the injection points 690 on
the mounting surface 242 and/or liquid fuel injected from the
atomizers 248, is directed by the air foils 610 to swirl in a
counter-clockwise direction (viewed from the combustion end)
through the annular swirl volume 255 (the volume between the swirl
vanes and the central hub). Within the annular swirl volume 255,
the continued swirling further mixes the fuel with the air.
[0077] The central hub 260 comprises an outer truncated cylindrical
conical surface centered on the central axis 200 of the nozzle to
minimize flow resistance to a circumferential flowing fuel-air
mixture from the main swirl vanes as it rises up into the central
flame holder 160. The central hub 260 forms a smooth surface rising
from the mounting surface 242 of the backplate 240 and sloping
concave inward to form a radial and axial support for the central
flame holder 160. Specifically, at its truncated upper reach, the
central hub 260 provides an outer annular support ledge 273 for the
central flame holder 160. The inner surface 263 of the central hub
260 defines a cavity 265 that accommodates a gas pilot nozzle 150
and includes an internal flow path for air to the gas pilot nozzle
150. The internal surface 263 of the central hub further includes
an inner annular mounting ledge 274 for the central flame holder
160.
[0078] The series of internal passages within the backplate
includes passages for outer main gas from the outer main gas
gallery in the endcover to the swirl vanes; for the inner main gas
gallery in the endcover to gas injectors on mounting surface of the
backplate; for liquid fuel from the liquid fuel delivery
penetrations in the endcover to atomizers on the mounting surface
of the backplate; and air passages from the circumferential outer
edge of the backplate to the central cavity for cooling and pilot
premix air to the radial nozzles center/core.
[0079] The internal passage 680 for inner main gas to the inner
main gas injector tips 695 on the mounting surface 242 of the
backplate 240 may include orifices 685 in each passage to control
gas flow rates to the gas injector tips 695. The outer
circumferential surface 257 of the backplate 240 includes a
plurality of radial feedholes 275 directed inward to the central
cavity 265 for feeding a flow of cooling air and pilot premix air
to the central cavity 265. The axial passages within the backplate
for outer main gas 270, inner main gas 680, and liquid fuel 247 are
situated in circumferential locations between the various radial
feedholes 275.
[0080] The central flame holder 160 may include a center hub 285, a
central cavity 278, a deswirler 280, and a plurality of v-gutters
290, an inner burner tube 300 and a support tower 295.
[0081] As airflow from between the main swirl vanes 250 is forced
into a rotational flow within the annular swirl volume 255, the
only exit path is downstream. About 30% of the fuel-air mixture
swirled in the main radial swirler 240 enters the central flame
holder 160. The central flame holder 160 includes the support tower
350 that sits atop the central hub 260 of the main radial swirler
240. The support tower 350 mates with the outer support ledge 273
and inner support ledge 274 of cylindrical support hub of the
central hub 260 to provide axial and radial support for the central
flame holder 160. Support arm 355 of the support tower 300 seats on
the outer support ledge 273 and the inner support ledge 274. A
central cavity 280 within the support tower 295 and the center hub
285 may accept the gas pilot nozzle 150.
[0082] Referring to FIG. 8, atop the support tower 295 sits the
deswirler 280 and the concentric hub-mounted, conical, v-gutter
flame-holder pack. The deswirler 280 includes a plurality of
annular compartments 345 between the center hub 285 and the inner
burner tube 295. The annular compartments 345 are open at upstream
entrance 347 and downstream exit 348 for the fuel-air mixture. A
radial vane 360 is provided between each individual annular
compartment 345, the radial vane extending from the center hub 285
to the inner burner tube 295. Each radial vane 360 curves from a
somewhat flat slope at the upstream entrance 347 to a steep slope
at the downstream exit 348 of the annular compartment 345. The flat
axial slope at the upstream entrance accommodates accepting the
swirling circumferential flow of the fuel-air mixture from the main
swirl vanes 250 of the main radial swirler 140. About 30% of
fuel-air mixture from the annular swirl volume 255 of the main
radial swirler flows 140 into the annular compartments 345 of the
deswirler 280. The changing slope of the radial vanes 360 redirects
the circumferential flow to an axial-oriented flow exiting each
individual annular compartment 360. The redirected axial flow
provides ventilation for the central recirculation zone (CRZ), as
previously described.
[0083] Referring to FIG. 3B and FIG. 8, an annular tip 380 of the
center hub 285 defines a plane surface I. An annular tip of the
inner burner tube 295 defines plane surface II. Plane surface I is
downstream of plane surface II. The radial vanes 360 of the
deswirler 280 at their downstream ends form a sloping edge between
the annular tip 380 of center hub 285 and the tip 385 of the inner
burner tube 300, rising at a slope of about 30%.
[0084] A v-gutter 290 is provided at the downstream end of each
radial vane 360. The v-gutter 290 comprises a v-shaped element 375
with the open end 376 facing downstream. A vertex 377 of the
v-shaped element 376 is attached to and extends through the annular
tip 380 of the center hub, along the downstream edge of radial wall
360 and through the tip 385 of inner burner tube 300.
[0085] An outer flame holder 170 comprises a generally cylindrical
outer burner tube 175, which flares at an upstream end to form a
annular seating surface for mating with the main swirler. The
cylindrical tube radially surrounds and extends towards the
combustion chamber beyond the central flame holder 160. The
downstream end 190 of the outer burner tube 175 is reinforced.
Ledge 195 provides a seating surface for engagement with the
conical dome 111 (FIG. 2) of the combustor. An annular seating
surface 180 of the outer burner tube 175 flares outward radially at
its upstream end. The seating surface 180 forms a roof over the
main swirlers 250 of the main radial swirler 140, thereby limiting
the exit path for the fuel-air mixture from the main radial swirler
140 to a downstream flow paths 402 and 405. The seating surface 180
may attach to the top of each airfoil for the main swirl vanes with
a plurality of bolts, one bolt for each tapped hole atop the
airfoil. The annular space formed between the inner burner tube 295
of the central flame holder 160 and the outer burner tube 175 of
outer flame holder 170 for flow 405 of the remaining 70% of the
fuel-air mixture from the main radial swirler 140 to the combustion
space.
[0086] FIG. 8 illustrates the central flame holder, gas pilot
annulus, and center cartridge for an embodiment of the inventive
large single radial nozzle. An air flow path for the remaining 5%
of incoming air from the inlet air plenum feeds to the central
cavity 260 of the nozzle 140 through a plurality of radial
penetrations 275 (12 in the embodiment) from the circumferential
edge 257 of the backplate 240.
[0087] As a means of achieving lightoff, combustor turndown, and
improving stability, a central gas pilot nozzle 150 is located
inside the conical flame-holder volume at the upstream,
smallest-diameter end. The gas pilot nozzle 150 provides a center
cartridge 155 which may include an igniter/flame detector and a
liquid gas pilot.
[0088] The roughly 5% of airflow to the radial nozzle that enters
through radial flowholes 275 in the circumferential surface 257 of
the backplate 240 for the main radial swirler 140 is split
internally. About 80% of this air flows forward through a air
supply annulus between the inner wall of the central hub central
cavity 265 and an outer surface 812 of an annular shell 810 of the
gas pilot nozzle 150 to an annular axial-swirled gas-pilot premixer
855. The remainder of the air passes through a plurality of radial
feed holes 875 in the annular shell 810 into the center cartridge
to be used for liquid-pilot atomization and cooling and purging of
the center-cartridge tip.
[0089] FIGS. 9A and 9B. illustrate the nozzle body of the gas pilot
nozzle for an embodiment of the inventive large single radial
nozzle.
[0090] The gas pilot nozzle 150 comprises a body 805 with annular
shell 810 that may be breech-loaded into the central cavity 203 of
the nozzle 140 through the endcover plate 205. The annular shell
810 includes aft flange 815 at its aft end with a plurality of bolt
holes 816 for mounting its forward surface 817 to the seating ledge
210 within the central cavity 203 of the endcover 205. The aft
flange 815 is also provided with a center hole 818 for insertion of
the center cartridge 155 and includes an elevated surface 819 on
the rear surface 820 about the center cavity incorporating tapped
holes 821 for bolting the center cartridge 155 to the gas pilot aft
flange 815. The aft flange 815 is also provided with a penetration
230 for connection to a pilot gas fuel supply for gas pilot
operation.
[0091] The gas pilot nozzle body 805 extends through the central
cavities 203, 265, 278 of the nozzle 120 and into the cylindrical
hub 370 of the central flame holder 160. The gas pilot annular
shell 810 tapers in steps from aft end to forward end. The annular
shell 810 includes lower shell 835, a tapered shell 840, a central
shell 845 and a tapered head 850.
[0092] An annular gas pilot airflow space 864 is also defined
between the inner wall 368 of the cylindrical hub 370 and the inner
wall 296 of support tower 295 with the outer surfaces 842, 847 of
the tapered shell 840 and the central shell 845. Air from inner
radial ends 277 of the central radial feedholes 275 in the
backplate 240 enters the gas pilot airflow space 864 and flows
forward axially to axial-swirled gas-pilot premixer 855.
[0093] The penetration 230 in aft flange 815 for pilot gas fuel
supplies internal pilot gas fuel cavities 862 in the annular shell
810. The internal pilot gas fuel cavities 842 within the lower
shell 835 supply pilot gas fuel to the annular pilot gas space 866
between the inner wall of annular shell 810 and outer surface 872
of center cartridge 155. The tapered head 850 extends in close
proximity to the cylindrical hub 370, thereby forming a gas pilot
annulus 825 between the outer surface 830 of the tapered annular
head 850 and the inner surface 368 of the cylindrical hub 370. A
plurality of pilot gas fuel holes 860 extend radially through the
annular shell at the upstream entrance between adjacent axial
mixing vanes 857 providing pilot gas fuel injection points. The
forward portion of the central shell 845 accommodates a plurality
of axial mixing vanes 857 in the general shape of airfoils on the
outer surface 847 for mixing gas pilot fuel and air moving
downstream in the airflow space 864, thereby constituting the
annular axial-swirled gas-pilot premixer 855.
[0094] The center-cartridge 155 includes cylindrical body 405
mounted on a rear flange 224. The center cartridge 155 is inserted
into the central cavity 203 of the gas pilot nozzle body 805 and
bolted through the rear flange 224 onto the raised rear surface
820. The rear flange 224 provides an axial penetration for
connections to an igniter and flame detector 236 and on its
circumferential surfaces, a radial penetration 232 for liquid pilot
fuel and a radial penetration 234 for an air The center cartridge
155 is aligned with central axis 200 of the nozzle.
[0095] FIG. 10 illustrates an axial tip-end section of the center
cartridge 155 for an embodiment of the inventive nozzle. The center
cartridge 155 is enclosed radially within a cartridge wall 872 and
at the downstream end by an end tip 885. The igniter 875 extends in
an axial direction from the center cartridge flange 224 to the end
tip 885. The liquid fuel pilot 880 extends from the center
cartridge flange 224 to the end tip 885. Air cavity 873 receives
air for use within the center cartridge. Air to the center
cartridge enters from the radial feedholes 275 in the backplate 240
and exits through holes 277 into the space 864 between the pilot
nozzle 150 and the inner surface 368 of the support tower 270. A
portion of the air entering the space 864 enters the center
cartridge 155 through the cartridge feedholes 870 filling the air
cavity around the igniter 875 and liquid fuel pilot 880 and
extending forward to a tip impingement shield 865. The tip
impingement shield 865 seals the upper end of the cavity and
includes a plurality of tip holes 867 (18 holes in the present
embodiment) Air from the tip impingement shield 865 is directed to
an annular air channel 876 at the downstream end of the igniter 875
to support ignition. Air from the tip impingement shield 865 is
also supplied to conical annulus 881 around a heat shield 882 on
the liquid fuel pilot 880. A plurality of offset blast holes 883
are provided through air blast atomizer shroud 884. The liquid
pilot fuel is provided through a cylindrical cavity 890 in liquid
pilot body 891. A truncated conical annulus of a spin chamber wall
892 at the tip defines a spin chamber 893 within for the liquid
fuel. An annular air gap 894 is provided around the liquid pilot
body 891 for thermal insulation. Around the annular air gap 894 is
provided an air assist annulus 895 connected to the air assist
supply in the flange 224 of the center cartridge 155. Within the
air assist annulus 895 is located an air-assist swirler 896
(embodiment includes 8 swirl vanes 897). The swirl vanes 897 impart
a swirling motion to the assist air being introduced to the spin
chamber 893.
[0096] FIG. 11 illustrates an alternate embodiment for the central
flame holder for the single large radial nozzle 900. Here the flame
holder 905 consists of a perforated cup 910. The perforated cup 910
consist of a plurality of holes 920 about the central axis 915,
through which 30% (approx) of the air fuel mixture 950 coming from
the annulus region passes, it partly mixes with the pilot air fuel
mixture 955 coming from the pilot nozzle 960 and partly burns at
the exit of these holes 920. The holes 920 are provided with
fillets 930 in order to minimize corner separation. A shroud 940 is
provided around the cup 910 to direct the flow into the cup. The
lower convex end 945 of the cup 920 is open and conformed to accept
the pilot air fuel mixture 955 from the gas pilot nozzle 960. Thus
in this case the heat release takes place in 3 stages. The first
stage is the pilot zone. The second stage is the air-fuel mixture
burning at the exit of these holes 920 and the third stage is the
flow bypassing the perforated cup 920.
[0097] The foregoing has described a single large radial nozzle for
a gas turbine combustor that provides major improvements in
operation over multi-nozzle designs. First, the intra-nozzle
combustion staging provided by the nozzles premixer design,
especially, the conical, de-swirled, v-gutter flame-holder in
conjunction with controllable outer main gas fuel injection paths
and inner main gas fuel-injection paths, is a unique aspect of this
design. This aspect allows multiple nozzles (per combustor) to be
replaced by one, resulting in major cost and part-count savings.
Second, combustion dynamics/oscillation abatement is created by the
smearing out of fuel transport times and heat release in the
chamber is a novel manner. This unique property additionally may
allow a broader range of fuels to be burned without needing to
modify or replace hardware. Finally, the combustor headend design
and the way the nozzle is integrated with the combustor dome in
creating an annular dome diffuser that recovers pressure while
convectively cooling the backside of the liner's dome without a
need for introduction of a separate cooling air source provides
increased simplicity with functionality.
[0098] Presently, the inventive nozzle has been sized for the GE
9FB heavy-duty industrial engine; however, it can be scaled up or
down in size to work for almost any combustor annular design (e.g.,
7H, 9H, 7FB, 7FA, 9FA, 6C, etc.). The design could be retrofitted
to an existing package, or it could be introduced as a new product
offering.
[0099] While only certain features of the invention have been
illustrated and described herein, many modifications and changes
will occur to those skilled in the art. It is, therefore, to be
understood that the appended claims are intended to cover all such
modifications and changes as fall within the true spirit of the
invention.
* * * * *