Note: Descriptions are shown in the official language in which they were submitted.
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METHOD AND APPARATUS FOR ACTIVELY CONTROLLING FUEL FLOW
TO A MIXER ASSEMBLY OF A GAS TURBINE ENGINE COMBUSTOR
BACKGROUND OF THE INVENTION
The present invention relates to a staged combustion system in which the
production
of undesirable combustion product components is minimized over the engine
operating regime and, more particularly, to a method and apparatus for
actively
controlling fuel flow to a mixer assembly having a pilot mixer with a primary
fuel
injector and secondary fuel injection ports.
Modern day emphasis on minimizing the production and discharge of gases that
contribute to smog and to other undesirable environmental conditions,
particularly
those gases that are emitted from internal combustion engines, have led to
different
gas turbine engine combustor designs that have been developed in an effort to
reduce
the production and discharge of such undesirable combustion product
components.
Other factors that influence combustor design are the desires of users of gas
turbine
engines for efficient, low cost operation, which translates into a need for
reduced fuel
consumption while at the same time maintaining or even increasing engine
output. As
a consequence, important design criteria for aircraft gas turbine engine
combustion
systems include provision for high combustion temperatures, in order to
provide high
thermal efficiency under a variety of engine operating conditions, as well as
the
minimization of undesirable combustion conditions that contribute to the
emission of
particulates, and to the emission of undesirable gases, and to the emission of
combustion products that are precursors to the formation of photochemical
smog.
Various governmental regulatory bodies have established emission limits for
acceptable levels of unburned hydrocarbons (HC), carbon monoxide (CO), and
oxides
of nitrogen (N0x), which have been identified as the primary contributors to
the
generation of undesirable atmospheric conditions. Therefore, different
combustor
designs have been developed to meet those criteria. For example, one way in
which
the problem of minimizing the emission of undesirable gas turbine engine
combustion
products has been attacked is the provision of staged combustion. In that
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arrangement, a combustor is provided in which a first stage burner is utilized
for low
speed and low power conditions to more closely control the character of the
combustion products. A combination of first stage and second stage burners is
provided for higher power outlet conditions while attempting to maintain the
combustion products within the emissions limits. It will be appreciated that
balancing
the operation of the first and second stage burners to allow efficient thermal
operation
of the engine, while simultaneously minimizing the production of undesirable
combustion products, is difficult to achieve. In that regard, operating at low
combustion temperatures to lower the emissions of NOx, can also result in
incomplete
or partially incomplete combustion, which can lead to the production of
excessive
amounts of HC and CO, in addition to producing lower power output and lower
thermal efficiency. High combustion temperature, on the other hand, although
improving thermal efficiency and lowering the amount of HC and CO, often
results in
a higher output of NOx.
Another way that has been proposed to minimize the production of those
undesirable
combustion product components is to provide for more effective intermixing of
the
injected fuel and the combustion air. In that regard, numerous mixer designs
have
been proposed over the years to improve the mixing of the fuel and air. In
this way,
burning occurs uniformly over the entire mixture and reduces the level of HC
and CO
that result from incomplete combustion. Even with improved mixing, however,
higher levels of undesirable NOx are formed under high power conditions when
the
flame temperatures are high.
One mixer design that has been utilized is known as a twin annular premixing
swirler
(TAPS), which is disclosed in the following U.S. Patents: 6,354,072;
6,363,726;
6,367,262; 6,381,964; 6,389,815; 6,418,726; 6,453,660; 6,484,489; and,
6,865,889. It
will be understood that the TAPS mixer assembly includes a pilot mixer which
is
supplied with fuel during the entire engine operating cycle and a main mixer
which is
supplied with fuel only during increased power conditions of the engine
operating
cycle. While improvements in the main mixer of the assembly during high power
conditions (i.e., take-off and climb) are disclosed in U.S. published patent
application
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Nos. US 2007-0028624 Al, US 2007-0028618 Al, and US 2007-0028617 Al,
modification of the pilot mixer is desired to improve operability across other
portions
of the engine's operating envelope (i.e., idle, approach and cruise) while
maintaining
combustion efficiency.
In order to provide increased functionality and flexibility, the pilot mixer
in a TAPS
type mixer assembly has been developed and is disclosed in a patent
application
entitled "Pilot Mixer For Mixer Assembly Of A Gas Turbine Engine Combustor
Having A Primary Fuel Injector And A Plurality Of Secondary Fuel Injection
Ports."
This patent application, having Canadian Serial No. 2,596,789 is owned by the
assignee of the present application. While the '789 application is concerned
with the
physical embodiments of the pilot mixer, it will be appreciated that an
apparatus and
method is desired which is able to actively control fuel flow to such pilot
mixer, as
well as the overall mixer assembly containing it.
It is well known that lean, premix combustion requires operation close to the
lean-
blow out boundary in order to minimize emissions. Therefore, it is desired
that the
onset of a lean blow out event be recognized so that operation of the
combustor can be
adjusted and lean blow out avoided. In addition, the mixing of air and fuel
must be
extremely effective to achieve low emissions. To enhance such mixing, pulsing
the
fuel to the injectors at a high frequency would also be desirable.
It has also been found that lean, premix combustion often results in high
dynamic
pressure levels in the combustor. The combustion dynamics is a result of
interaction
between heat release from combusting the fuel-air mixture and pressure
oscillations in
the chamber. Such dynamic pressures may result in high cycle fatigue and can
damage combustor parts. While the effects of dynamic pressures on the
combustor
have been countered previously, this has generally involved the provision of
high
bandwidth fuel or air actuation to reduce the pressure levels associated with
acoustic
modes of the combustor.
Thus, there is a need to provide a gas turbine engine combustor in which the
production of undesirable combustion product components is minimized over a
wide
range of engine operating conditions. Accordingly, it is desired that the
pilot mixer of
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a nested combustor arrangement be modified to include a primary fuel injector
and a
plurality of secondary fuel injection ports. It is also desired that an active
control
system and process be provided which enhances operation of such mixer assembly
by
identifying and countering the onset of a lean blow out condition, as well as
an
unacceptable level of dynamic pressure experienced in the combustor.
BRIEF SUMMARY OF THE INVENTION
In a first exemplary embodiment of the invention, an apparatus for actively
controlling
fuel flow from a fuel pump to a mixer assembly of a gas turbine engine
combustor is
disclosed, where the mixer assembly includes a pilot mixer and a main mixer.
The
pilot mixer further includes an annular pilot housing having a hollow
interior, a
primary fuel injector mounted in the pilot housing and adapted for dispensing
droplets
of fuel to the hollow interior of the pilot housing, and a plurality of axial
swirlers
positioned upstream from the primary fuel injector. The fuel flow control
apparatus
further includes: at least one sensor for detecting dynamic pressure in the
combustor;
a fuel nozzle; and, a system for actively controlling fuel flow supplied to
the pilot
mixer and the main mixer of the mixer assembly by the fuel nozzle. The fuel
nozzle
further includes: a feed strip with a plurality of circuits for providing fuel
to the pilot
mixer and the main mixer; and, a plurality of valves associated with the fuel
nozzle
and in flow communication with the feed strip thereof The control system
activates
the valves in accordance with signals received from the pressure sensor.
In a second exemplary embodiment of the invention, an apparatus for actively
controlling fuel flow from a fuel pump to a mixer assembly of a gas turbine
engine
combustor is disclosed, where the mixer assembly includes a pilot mixer and a
main
mixer. The pilot mixer further includes an annular pilot housing having a
hollow
interior, a primary fuel injector mounted in the pilot housing and adapted for
dispensing droplets of fuel to the hollow interior of the pilot housing, a
plurality of
axial swirlers positioned upstream from the primary fuel injector, and a
plurality of
secondary fuel injection ports for introducing fuel into the hollow interior
of the pilot
housing. The fuel flow control apparatus further includes: at least one sensor
for
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detecting dynamic pressure in the combustor; a fuel nozzle; and, a system for
actively
controlling fuel flow supplied to the pilot mixer and the main mixer of the
mixer
assembly by the fuel nozzle. The fuel nozzle further includes: a feed strip
with a
plurality of circuits for providing fuel to the primary fuel injector of the
pilot mixer,
the secondary fuel injection ports of the pilot mixer, and the main mixer;
and, a
plurality of valves associated with the fuel nozzle and in flow communication
with the
feed strip thereof. The control system activates the valves in accordance with
signals
received from the pressure sensor.
In a third exemplary embodiment of the invention, a method of actively
controlling
fuel flow from a fuel pump to a mixer assembly of a gas turbine engine
combustor is
disclosed, the mixer assembly including a pilot mixer and a main mixer,
wherein the
pilot mixer further includes an annular pilot housing having a hollow interior
and a
primary fuel injector mounted in the pilot housing and adapted for dispensing
droplets
of fuel to the hollow interior of the pilot housing. The method includes the
following
steps: continuously sensing dynamic pressure in a combustion chamber of the
combustor; determining whether an amplitude of the sensed dynamic pressure in
the
combustion chamber is greater than a predetermined amount; and, signaling a
fuel
nozzle to provide fuel in a specified manner to the pilot mixer when the
pressure
amplitude is greater than the predetermined amount.
In a fourth exemplary embodiment of the invention, a method of actively
controlling
fuel flow from a fuel pump to a mixer assembly of a gas turbine engine
combustor is
disclosed, the mixer assembly including a pilot mixer and a main mixer,
wherein the
pilot mixer further includes an annular pilot housing having a hollow
interior, a
primary fuel injector mounted in the pilot housing and adapted for dispensing
droplets
of fuel to the hollow interior of the pilot housing, and a plurality of
secondary fuel
injection ports for introducing fuel into the hollow interior of the pilot
housing. The
method includes the following steps: continuously sensing dynamic pressure in
a
combustion chamber of the combustor; determining whether an amplitude of the
sensed dynamic pressure in the combustion chamber is greater than a
predetermined
amount; and, signaling a fuel nozzle to provide fuel in a specified manner to
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secondary fuel injection ports of the pilot mixer when the pressure amplitude
is
greater than the predetermined amount.
In a fifth exemplary embodiment of the invention, a method of actively
controlling
fuel flow from a fuel pump to a mixer assembly of a gas turbine engine
combustor is
disclosed, the mixer assembly including a pilot mixer and a main mixer,
wherein the
pilot mixer further includes an annular pilot housing having a hollow interior
and a
primary fuel injector mounted in the pilot housing and adapted for dispensing
droplets
of fuel to the hollow interior of the pilot housing. The method includes the
following
steps: continuously sensing dynamic pressure in a combustion chamber of the
combustor; determining whether a frequency of the sensed dynamic pressure in
the
combustion chamber is within a predetermined range; and, signaling a fuel
nozzle to
provide fuel in a specified manner to the pilot mixer when the pressure
frequency is
within the predetermined range.
In a sixth exemplary embodiment of the invention, a method of actively
controlling
fuel flow from a fuel pump to a mixer assembly of a gas turbine engine
combustor is
disclosed, the mixer assembly including a pilot mixer and a main mixer,
wherein the
pilot mixer further includes an annular pilot housing having a hollow
interior, a
primary fuel injector mounted in the pilot housing and adapted for dispensing
droplets
of fuel to the hollow interior of the pilot housing, and a plurality of
secondary fuel
injection ports for introducing fuel into the hollow interior of the pilot
housing. The
method includes the following steps: continuously sensing dynamic pressure in
a
combustion chamber of the combustor; determining whether a frequency of the
sensed
dynamic pressure in the combustion chamber is within a predetermined range;
and,
signaling a fuel nozzle to provide fuel in a specified manner to the secondary
fuel
injection ports of the pilot mixer when the pressure frequency is within the
predetermined range.
In a seventh exemplary embodiment of the invention, a method of actively
controlling
fuel flow from a fuel pump to a mixer assembly of a gas turbine engine
combustor
during a plurality of operational stages is disclosed, the mixer assembly
including a
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pilot mixer and a main mixer, wherein the pilot mixer further includes an
annular pilot
housing having a hollow interior, a primary fuel injector mounted in the pilot
housing
and adapted for dispensing droplets of fuel to the hollow interior of the
pilot housing,
and a plurality of secondary fuel injection ports for introducing fuel into
the hollow
interior of the pilot housing. The method includes the following steps:
supplying fuel
only to the primary fuel injector and the secondary fuel injection ports of
the pilot
mixer during a first fueling mode; supplying fuel to the pilot mixer and the
main mixer
in a first specified amount during a second fueling mode; and, supplying fuel
to the pilot
mixer and the main mixer in a second specified amount during a third fueling
mode.
BRIEF DESCRIPTION OF THE DRAWINGS
FIG. 1 is a diagrammatic view of a high bypass turbofan gas turbine engine;
FIG. 2 is a longitudinal, cross-sectional view of a gas turbine engine
combustor having
a staged arrangement;
FIG. 3 is an enlarged, cross-sectional view of the mixer assembly depicted in
Fig. 2;
FIG. 4 is a cross-sectional view of a fuel nozzle assembly and the mixer
assembly
depicted in Figs. 2 and 3;
FIG. 5 is a block diagram of a system for providing fuel flow to the mixer
assembly
depicted in Figs. 2 and 3;
FIG. 6 is a schematic diagram of a system for actively controlling fuel flow
through
the fuel nozzle assembly depicted in Fig. 4;
FIG. 7 is a flow diagram depicting operational steps involved in a process for
actively
controlling fuel flow through the fuel nozzle assembly depicted in Fig. 4 to
the mixer
assembly depicted in Figs. 2 and 3; and,
FIG. 8 is a diagrammatic view of how fuel flow is provided to the mixer
assembly
depicted in Figs. 2 and 3 during specified stages of engine operation.
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DETAILED DESCRIPTION OF THE INVENTION
Referring now to the drawings in detail, wherein identical numerals indicate
the same
elements throughout the figures, Fig. 1 depicts in diagrammatic form an
exemplary
gas turbine engine 10 (high bypass type) utilized with aircraft having a
longitudinal or
axial centerline axis 12 therethrough for reference purposes. Engine 10
preferably
includes a core gas turbine engine generally identified by numeral 14 and a
fan section
16 positioned upstream thereof. Core engine 14 typically includes a generally
tubular
outer casing 18 that defines an annular inlet 20. Outer casing 18 further
encloses and
supports a booster compressor 22 for raising the pressure of the air that
enters core
engine 14 to a first pressure level. A high pressure, multi-stage, axial-flow
compressor 24 receives pressurized air from booster 22 and further increases
the
pressure of the air. The pressurized air flows to a combustor 26, where fuel
is injected
into the pressurized air stream to raise the temperature and energy level of
the
pressurized air. The high energy combustion products flow from combustor 26 to
a
first (high pressure) turbine 28 for driving high pressure compressor 24
through a first
(high pressure) drive shaft 30, and then to a second (low pressure) turbine 32
for
driving booster compressor 22 and fan section 16 through a second (low
pressure)
drive shaft 34 that is coaxial with first drive shaft 30. After driving each
of turbines
28 and 32, the combustion products leave core engine 14 through an exhaust
nozzle
36 to provide propulsive jet thrust.
Fan section 16 includes a rotatable, axial-flow fan rotor 38 that is
surrounded by an
annular fan casing 40. It will be appreciated that fan casing 40 is supported
from core
engine 14 by a plurality of substantially radially-extending,
circumferentially-spaced
outlet guide vanes 42. In this way, fan casing 40 encloses fan rotor 38 and
fan rotor
blades 44. Downstream section 46 of fan casing 40 extends over an outer
portion of
core engine 14 to define a secondary, or bypass, airflow conduit 48 that
provides
additional propulsive jet thrust.
From a flow standpoint, it will be appreciated that an initial air flow,
represented by
arrow 50, enters gas turbine engine 10 through an inlet 52 to fan casing 40.
Air flow
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50 passes through fan blades 44 and splits into a first compressed air flow
(represented
by arrow 54) that moves through conduit 48 and a second compressed air flow
(represented by arrow 56) which enters booster compressor 22. The pressure of
second compressed air flow 56 is increased and enters high pressure compressor
24, as
represented by arrow 58. After mixing with fuel and being combusted in
combustor
26, combustion products 60 exit combustor 26 and flow through first turbine
28.
Combustion products 60 then flow through second turbine 32 and exit exhaust
nozzle
36 to provide thrust for gas turbine engine 10.
As best seen in Fig. 2, combustor 26 includes an annular combustion chamber 62
that
is coaxial with longitudinal axis 12, as well as an inlet 64 and an outlet 66.
As noted
above, combustor 26 receives an annular stream of pressurized air from a high
pressure compressor discharge outlet 69. A portion of this compressor
discharge air
flows into a mixing assembly 67, where fuel is also injected from a fuel
nozzle 68 to
mix with the air and form a fuel-air mixture that is provided to combustion
chamber
62 for combustion. Ignition of the fuel-air mixture is accomplished by a
suitable
igniter (not shown), and the resulting combustion gases 60 flow in an axial
direction
toward and into an annular, first stage turbine nozzle 72. Nozzle 72 is
defined by an
annular flow channel that includes a plurality of radially-extending,
circularly-spaced
nozzle vanes 74 that turn the gases so that they flow angularly and impinge
upon the
first stage turbine blades of first turbine 28. As shown in Fig. 1, first
turbine 28
preferably rotates high pressure compressor 24 via first drive shaft 30. Low
pressure
turbine 32 preferably drives booster compressor 24 and fan rotor 38 via second
drive
shaft 34.
Combustion chamber 62 is housed within engine outer casing 18 and is defined
by an
annular combustor outer liner 76 and a radially-inwardly positioned annular
combustor inner liner 78. The arrows in Fig. 2 show the directions in which
compressor discharge air flows within combustor 26. As shown, part of the air
flows
over the outermost surface of outer liner 76, part flows into combustion
chamber 62,
and part flows over the innermost surface of inner liner 78.
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Contrary to previous designs, it is preferred that outer and inner liners 76
and 78,
respectively, not be provided with a plurality of dilution openings to allow
additional
air to enter combustion chamber 62 for completion of the combustion process
before
the combustion products enter turbine nozzle 72. This is in accordance with a
patent
application entitled "High Pressure Gas Turbine Engine Having Reduced
Emissions"
and having U.S. published patent application US 2007-0028595 Al, which is also
owned by the assignee of the present invention. It will be understood,
however, that
outer liner 76 and inner liner 78 preferably include a plurality of smaller,
circularly-
spaced cooling air apertures (not shown) for allowing some of the air that
flows along
the outermost surfaces thereof to flow into the interior of combustion chamber
62.
Those inwardly-directed air flows pass along the inner surfaces of outer and
inner
liners 76 and 78 that face the interior of combustion chamber 62 so that a
film of
cooling air is provided therealong.
It will be understood that a plurality of axially-extending mixing assemblies
67 are
disposed in a circular array at the upstream end of combustor 26 and extend
into inlet
64 of annular combustion chamber 62. It will be seen that an annular dome
plate 80
extends inwardly and forwardly to define an upstream end of combustion chamber
62
and has a plurality of circumferentially spaced openings formed therein for
receiving
mixing assemblies 67. For their part, upstream portions of each of inner and
outer
liners 76 and 78, respectively, are spaced from each other in a radial
direction and
define an outer cowl 82 and an inner cowl 84. The spacing between the
forwardmost
ends of outer and inner cowls 82 and 84 defines combustion chamber inlet 64 to
provide an opening to allow compressor discharge air to enter combustion
chamber 62.
A mixing assembly 100 in accordance with one embodiment of the present
invention
is shown in FIG. 3. Mixing assembly 100 preferably includes a pilot mixer 102,
a
main mixer 104, and a cavity 106 positioned therebetween. More specifically,
it will
be seen that pilot mixer 102 preferably includes an annular pilot housing 108
having a
hollow interior, as well as a primary fuel injector 110 mounted in housing 108
and
adapted for dispensing droplets of fuel to the hollow interior of pilot
housing 108.
Further, pilot mixer 102 preferably includes a first swirler 112 located at a
radially
inner position adjacent primary fuel injector 110, a second swirler 114
located at a
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radially outer position from first swirler 112, and a splitter 116 positioned
therebetween. As shown, splitter 116 extends downstream of primary fuel
injector
110 to form a venturi 118 at a downstream portion. It will be understood that
first and
second pilot swirlers 112 and 114 are generally oriented parallel to a
centerline axis
120 through mixing assembly 100 and include a plurality of vanes for swirling
air
traveling therethrough. Fuel and air are provided to pilot mixer 102 at all
times during
the engine operating cycle so that a primary combustion zone 122 is produced
within a
central portion of combustion chamber 62 (see Fig. 2).
Main mixer 104 further includes an annular main housing 124 radially
surrounding
pilot housing 108 and defining an annular cavity 126, a plurality of fuel
injection ports
128 which introduce fuel into annular cavity 126, and a swirler arrangement
identified
generally by numeral 130. Swirler arrangement 130 may be configured in any of
several ways, as seen in U.S. published patent application entitled "Mixer
Assembly
For Combustor Of A Gas Turbine Engine Having A Plurality Of Counter-Rotating
Swirlers" having No. US 2007-0028624 Al and U.S. published patent application
entitled "Swirler Arrangement For Mixer Assembly Of A Gas Turbine Engine
Combustor Having Shaped Passages" having No. US 2007-0017224 Al, both of
which are assigned to the owner of the present invention. It will be seen in
Fig. 3,
however, that swirler arrangement 130 preferably includes at least a first
swirler 144
positioned upstream from fuel injection ports 128. As shown, first swirler 144
is
preferably oriented substantially radially to centerline axis 120 through
mixer
assembly 100. It will be noted that first swirler 144 includes a plurality of
vanes 150
for swirling the air flowing therebetween. Since vanes 150 are substantially
uniformly
spaced circumferentially, a plurality of substantially uniform passages are
defined
between adjacent vanes 150. It will further be understood that swirler 144 may
include vanes having different configurations so as to shape the passages in a
desirable
manner, as disclosed in the '244 patent application identified hereinabove.
Swirler arrangement 130 also is shown as including a second swirler 146
positioned
upstream from fuel injection ports 128 and preferably oriented substantially
parallel to
centerline axis 120. Second swirler 146 further includes a plurality of vanes
152 for
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swirling the air flowing therebetween. Although vanes 152 are shown as being
substantially uniformly spaced circumferentially, thereby defining a plurality
of
substantially uniform passages therebetween, such vanes 152 may also have
different
configurations so as to shape the passages in a desirable manner.
Cavity 106, as stated above, is located between pilot mixer 102 and main mixer
104
and contains a first fuel manifold 107 in flow communication with a fuel
supply. In
particular, a centerbody outer shell 140 forms an outer surface and an aft
surface of
cavity 106, with pilot housing 108 providing an inner surface thereof. Fuel
injection
ports 128 are in flow communication with fuel manifold 107 and spaced
circumferentially around centerbody outer shell 140. As seen in Fig. 3, fuel
injection
ports 128 are preferably positioned so that fuel is provided in an upstream
end of
annular cavity 126.
When fuel is provided to main mixer 104, an annular, secondary combustion zone
198
is provided in combustion chamber 62 that is radially outwardly spaced from
and
concentrically surrounds primary combustion zone 122. Depending upon the size
of
gas turbine engine 10, as many as twenty or so mixer assemblies 100 can be
disposed
in a circular array at inlet 64 of combustion chamber 62.
As seen in Fig. 3, pilot mixer 102 also preferably includes a plurality of
spaced
secondary fuel injection ports 134, whereby fuel is also introduced into
hollow interior
of pilot housing 108. It will be appreciated that secondary fuel injection
ports 134 are
preferably spaced circumferentially about pilot housing 108 within a
designated plane
136 intersecting centerline axis 120 through mixing assembly 100. While plane
136,
in which secondary fuel injection ports 134 lie, is shown as being located in
a flared
portion 138 of pilot housing 108 downstream of splitter 116, it will be
understood that
a plane containing such secondary fuel injection ports 134 may be located at
approximately a downstream end of splitter 116 or even upstream thereof.
Indeed, the
axial length of splitter 116 may be altered so that its relationship with the
location of
secondary fuel injection ports 134 could change.
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Similarly, plane 136 is depicted as being oriented substantially perpendicular
to
centerline axis 120, but secondary fuel injection ports 134 may be positioned
so that
plane 136 is skewed so as to be angled either upstream or downstream as
desired.
Further, regardless of the axial position or orientation of plane 136
containing
secondary fuel injection ports 134, each such secondary fuel injection port
134 may
individually be oriented substantially perpendicular to centerline axis 120,
oriented
upstream at an acute angle, or oriented downstream at an obtuse angle.
It will further be seen that secondary fuel injection ports 134 of pilot mixer
102
preferably are in flow communication with a second fuel manifold 109, which
also is
preferably located within cavity 106. Fuel is typically injected into the
hollow portion
of pilot housing 108 by secondary fuel injection ports 134 upon the occurrence
of a
specified event (e.g., a designated cycle point for gas turbine engine 10,
when
compressor discharge air 58 is a designated temperature, etc.). Depending upon
the
requirements of a specific condition, fuel is injected through secondary fuel
injection
ports 134 at a rate greater than, less than or substantially the same as fuel
injected
through primary fuel injector 110. Of course, this presumes that fuel will be
provided
by primary fuel injector 110 at all times, but there may be occasions when it
is preferable
to provide fuel to pilot mixer 102 only through secondary fuel injection ports
134.
In this way, pilot mixer 102 has greater flexibility during operation across
the lower
power conditions (i.e., idle, approach and cruise). In particular, it will be
appreciated
that pilot mixer 102 is able to power gas turbine engine 10 up to
approximately 30%
of maximum thrust when fuel is provided solely to primary fuel injector 110.
By
comparison, pilot mixer 102 is able to power gas turbine engine 10 up to
approximately 70% of maximum thrust when fuel is provided to secondary fuel
injection ports 134 as well.
In order to promote the desired fuel spray into the hollow interior of pilot
housing 108,
it is preferred that a passage 142 surround each secondary fuel injection port
134 of
pilot mixer 102. Each passage 142 is in flow communication with compressed air
via
a supply 154 provided in cavity 106. This air is provided to facilitate
injection of the
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fuel spray into pilot housing 108 instead of being forced along an inner
surface 156
thereof. This may further be enhanced by providing a swirler 158 within each
passage
142 which provides a swirl to the air injected around the fuel spray.
It is also preferred that vanes of outer pilot swirler 114 be configured so
that air
passing therethrough is directed at least somewhat toward inner surface 156 of
pilot
housing 108. In this way, such air is better able to interact with fuel
provided by
secondary fuel injection ports 134. Accordingly, such vanes are preferably
angled at
approximately 30 to about 60 with respect to centerline axis 120. In this
way, a flare
angle 160 of pilot housing 108 is approximated.
Considering the addition of secondary fuel injection ports 134 in pilot mixer
102, it
will be appreciated that the flow rate of air therethrough is preferably
maintained at a
rate of approximately 10% to approximately 30%. Further, such secondary
injection
ports 134 assist in reducing the emissions produced by mixer assembly 100
during the
operation of gas turbine engine 10. In particular, combustor 26 is able to
operate only
with fuel being supplied to pilot mixer 102 for a greater time period. Also,
it has been
found that providing more fuel at a radially outer location of pilot mixer 102
is desirable.
It will further be seen in Figs. 4-7 that an apparatus and method for
controlling fuel
flow to mixer assembly 100 is provided. With respect to fuel nozzle 68, it
will be
appreciated that it is configured similar to that shown and described in U.S.
Patent
6,955,040 to Myers, Jr. et al. More specifically, it will be seen that fuel
nozzle 68
includes a housing 174 located at an outer radial location which contains a
plurality of
valves, a nozzle support 176 which extends between valve housing 174 and mixer
assembly 100, and a macrolaminate feed strip 178 positioned within nozzle
support
176. Feed strip 178 further includes a first circuit 180 for supplying fuel to
a fuel tube
132 (which is in flow communication with primary fuel injector 110 of pilot
mixer
102), a second circuit 182 for supplying fuel to fuel manifold 109 (which is
in flow
communication with secondary fuel injection ports 134 of pilot mixer 102), and
a third
circuit 183 for supplying fuel to fuel manifold 107 (which is in flow
communication
with fuel injection ports 128 of main mixer 104).
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In order to better understand the manner in which fuel is supplied to mixer
assembly
100, a block diagram of an overall fuel flow control system 200 is depicted in
Fig. 5.
As seen therein, system 200 includes a fuel pump 202, whereby a fuel supply
(not
shown) in flow communication therewith provides fuel to each fuel nozzle 68
positioned around annular combustor 26. A fuel nozzle control 204 is provided
for
each fuel nozzle 68 in order to generally control the valves within housing
174 and
therefore the amount of fuel provided by circuits 180, 182 and 183. Fuel
nozzle
control 204 interfaces with fuel pump 202 and receives signals 208 from a full
authority digital engine control (FADEC) 206 to coordinate the proper fueling
mode
of pilot and main mixers 102 and 104 depending upon the current stage of
operational
cycle for gas turbine engine 10. This will be explained in more detail herein
with
respect to Fig. 8.
It will be appreciated that staging valves 184, 186 and 188, which are
associated with
circuits 180, 182 and 183, respectively, are activated according to a signal
210
provided by fuel nozzle control 204. Fuel is then permitted to flow through
first
circuit 180, second circuit 182, and third circuit 183 within feed strip 178
of each fuel
nozzle 68 according to the positioning of staging valves 184, 186 and 188. In
this
way, fuel is either provided in the desired amount to primary fuel injector
110 of pilot
mixer 102, secondary fuel injection ports 134 of pilot mixer 102, and fuel
injection
ports 128 of main mixer 104 of each mixer assembly 100.
In order to pulse the fuel in first, second, and/or third circuits 180, 182
and 183, a
second separate control signal 212 from engine control 206 is provided to a
pulsing
valve 185, a pulsing valve 187 and/or a pulsing valve 189, respectively, of
each fuel
nozzle 68. It will be noted that pulsing valves 185, 187 and 189 are located
within a
pulsing valve housing 191 (see Fig. 4). Among other various readings, signals
and
measurements received by engine control 206, a signal 216 is also provided
thereto by
at least one pressure sensor 218 located adjacent to outer liner 76 of
combustor 26 (see
Fig. 2). Pressure sensor 218 senses a frequency and an amplitude for the
pressure
within combustion chamber 62 and imparts this information to engine control
206 via
signal 216. Pressure sensor 218 is capable of withstanding high temperatures
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experienced in combustion chamber 62. Accordingly, an exemplary pressure
sensor is
a diaphragm type of transducer, where the displacement of the diaphragm is
proportional to the dynamic component of the input pressure signal. While only
one
pressure sensor 218 is depicted in Figs. 2 and 5, it is preferred that a
plurality of
pressures sensors 218 be equally spaced circumferentially around outer liner
76 in
order to detect dynamic pressure of combustion chamber 62 in a more localized
region. Accordingly, only those mixer assemblies located adjacent a region of
combustion chamber 62 experiencing dynamic instability are modulated
More specifically, Fig. 6 depicts a schematic diagram indicating the flow of
fuel from
fuel pump 202 to circuits 180, 182 and 183. It will be seen that fuel pump
202, which
includes both a booster pump 220 and a main pump 222, receives fuel from an
inlet
224. Fuel pump 202 sends fuel through a line 226 to a metering valve 228,
where the
pressure is controlled. In order to maintain a desired pressure for the fuel
entering
main pump 222, a bypass circuit 230 is in flow communication with line 226.
Bypass
circuit 230 includes a bypass line 232 with a bypass valve 234 therein for
controlling
flow back to main pump 222 via a bypass input line 236. It will also be noted
that fuel
nozzle control 204 taps into line 226 upstream of metering valve 228 via line
238 so
that it receives a high pressure source to modulate.
Upon exiting metering valve 228, line 240 splits first into a fuel supply line
242 that
provides fuel to a fuel supply manifold 244, which in turn supplies fuel to
valve
housing 174 of each fuel nozzle 68. A line 246 also in flow communication with
line
240 is connected to fuel nozzle control 202, which enables it to determine a
differential pressure control of the pressure control nozzle (DPCPFN) and a
torque
motor current of the pressure control nozzle (TMCPFN). From this information,
a
fuel signal circuit 248 from fuel nozzle control 202 controls the activation
of staging
valves 184, 186 and 188. More specifically, fuel signal circuit 248 includes
signal
210, also understood herein to be a pressure control pressure off the fuel
nozzle
(PCPFN), to a fuel signal manifold 250, whereupon fuel signal manifold 250
then
provides a signal 252 to each valve housing 174. It will be appreciated that
staging
valves 184, 186 and 188 will generally be activated according to signal 252 so
that the
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desired amount of fuel provided via fuel supply manifold 244 is passed to the
respective circuit of pilot mixer 102 (i.e., first and second circuit 180 and
182) and
main mixer 104 (i.e., third circuit 183).
A signal fuel return line 254 extends from each valve housing 174 so as to be
in flow
communication with fuel pump inlet 224. A sink line 256 from fuel nozzle
control
also connects to signal fuel return line 254.
It will be further seen in Fig. 6 that fuel nozzle control 202 receives signal
208 from
engine control 206. Under certain specified conditions, signal 208 instructs
fuel
nozzle control 202 to alter the distribution of fuel to circuits 180, 182 and
183 by
activating staging valves 184, 186 and 188 in a different manner. This occurs
when
the amplitude of a dynamic pressure instability is detected in combustion
chamber 62
above a predetermined level by one or more pressure sensors 218. While this
predetermined pressure amplitude level may vary or be conditioned upon other
engine
factors, it generally will be set at a level where integrity of the combustor
hardware is
maintained (e.g., approximately 0.5 psi peak to peak).
Besides altering the fuel split between circuits 180, 182, and 183, engine
control 206
may respond to such pressure instability by causing fuel to be pulsed through
one or
more of pulsing valves 185, 187 and/or 189. Pulsing of fuel through secondary
fuel
injection ports 134 of pilot mixer 102 in at least one mixing assembly 100
located near
the occurrence of the dynamic pressure instability, via pulsing valve 187, is
typically
preferred. It has been found that pulsing the fuel with an amplitude and
frequency
opposite that of the pressure dynamic reduces the pressure instability in that
location
of combustion chamber 62. Alternatively, pulsing of the fuel may be done at an
amplitude and frequency which is a sub-harmonic of the dynamic pressure on the
combustion chamber. Pulsing fuel in this way would be at a lower bandwidth,
which
would reduce the stress on pulsing valve 187 and increase the life thereof. By
utilizing a closed loop system of detecting pressure instabilities and then
offsetting
them through the pulsing of fuel in this way, the problem is attacked
continuously
until the dynamic pressure instability is below the predetermined level.
Although fuel
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could alternatively be pulsed through primary fuel injector 110 of pilot mixer
102 via
and/or fuel injection ports 128 of main mixer 104 to offset dynamic pressure
instabilities in combustion chamber 62, such as the case when pilot mixer 102
does
not include secondary fuel injection ports 134, it will be appreciated that
pulsing fuel
flow to secondary injection ports 134 has a minimal effect on the fuel/air
mixture
within mixing assembly 100.
It has also been found that a frequency signal from pressure sensors 218
within a
specified range is indicative of an incipient lean blowout condition for
combustor 26.
This signal range is approximately 40 Hertz to approximately 50 Hertz and is
able to
predict the oncoming condition as opposed to merely detecting it. Accordingly,
an
override signal 214 is preferably provided by engine control 206 to valve
housing 174
so that additional fuel can be supplied to mixer assembly 100. Preferably,
override
signal 214 involves the activation of valve 186, whereby additional fuel is
injected
into pilot mixer 102 by means of secondary fuel injection ports 134. The fuel
split
between pilot mixer 102 and main mixer 104 may also be altered by increasing
the
amount of fuel provided to primary fuel injector 110 (e.g., when pilot mixer
102 does
not include secondary fuel injection ports 134).
Thus, it will be appreciated that modifying the fuel split between pilot mixer
102 and
main mixer 104, and even between primary fuel injector 110 and secondary fuel
injection ports 134, effectively counters the dynamic pressure instabilities
in
combustion chamber 62 and an incipient lean blow out condition for combustor
26.
Likewise, pulsing fuel in primary fuel injector 110, secondary fuel injection
ports 134,
and/or fuel injection ports 128 is effective for the same purposes.
It will also be understood that control system 200 is also effective for
controlling the
pressure dynamics in combustor 26 when actions therein are initiated
intentionally.
For example, it may be desirable in certain instances (e.g., to improve the
mixing of
fuel and air during fuel rich conditions) to pulse the fuel provided to mixer
100. Such
pulsing of fuel in and of itself may create pressure dynamics which need to be
maintained within acceptable limits. Detection and control of such pressure
dynamics
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by means of pressure sensors 218 and engine control 206 may cause the pulsing
of
fuel to be modified accordingly.
In conjunction with the physical embodiments of mixer assembly 100 and fuel
flow
control system 200, it will be understood from the flow diagram in Fig. 7 that
a
method of actively controlling fuel flow to mixer assembly 100 is also
presented.
More specifically, such method includes the following steps: sensing dynamic
pressure (frequency and amplitude) in combustion chamber 62 of combustor 26
via
pressure sensors 218 (box 260); providing signal 216 containing frequency and
amplitude information of such pressure to engine control 206 (box 262); and,
determining whether the frequency component of signal 216 is within a
specified
range indicative of incipient lean blow out (comparator box 264). If the
frequency
component of the pressure signal 216 is within such specified frequency range,
then
engine control 206 provides signal 214 to valve housing 174 to override the
current
status of staging valves 184, 186 and 188 to inject additional fuel into pilot
mixer 102
(box 266). Afterward, the dynamic pressure in combustion chamber 62 continues
to
be sensed as represented by a feedback loop 267 to box 260.
Should the frequency component of signal 216 not be within the specified
frequency
range, then the next step in the process is determining whether an amplitude
component of signal 216 is greater than the predetermined level indicative of
a
dynamic instability (comparator box 268). If this is found to be so, then
engine
control 206 provides signal 212 to activate pulsing valve 187 (and/or pulsing
valves
185 and 189) and thereby modulate pilot flow at a frequency and amplitude
which
abates the dynamic instability (box 270). Thereafter, the dynamic pressure in
combustion chamber 62 continues to be sensed as represented by a feedback loop
272.
Should the amplitude component of signal 216 be less than the predetermined
level,
the system likewise returns to sensing the dynamic pressure in combustion
chamber
62 as shown by feedback 274 connecting to feedback loop 272.
Fig. 8 further illustrates a staging diagram for mixer assembly 100, whereby
the
relative amount of fuel provided to pilot mixer 102 and main mixer 104 is
provided
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for various points in the cycle of engine 10 (i.e., to obtain certain
temperature ranges
for combustor 26). Because pilot mixer 102 includes both primary fuel injector
110
and secondary fuel injection ports 134, it has been found that engine 10 is
able to
operate at an extended temperature range when only providing fuel thereto.
This also
enables fuel nozzle control 204 to eliminate a separate fueling mode (i.e.,
60% pilot
mixer/40% main mixer) which has been utilized previously. As seen in a bar 275
in
Fig. 8, the first fueling mode involves 100% of the fuel being provided to
pilot mixer
102 to obtain a combustor temperature range of approximately 200 F to about
800 F.
Bar 275 further depicts that a first cross-hatched portion 276 thereof is
attributed to
fuel being provided only to primary fuel injector 110 (i.e., to obtain a
combustor
temperature range of approximately 200 F to approximately 500 F) and a second
cross-hatched portion 278 represents fuel being provided to both primary fuel
injector
110 and secondary fuel injection ports 134 (to obtain a combustor temperature
range
of approximately 500 F to approximately 800 F). This first stage is considered
to be
the range of normal operation for optimum performance of combustor 26 when
pilot
mixer 102 only is fueled. Thus, this first fueling mode is typically used for
idle, taxi
and approach portions of engine operation.
It has been found that a fuel pump limit 281 for the first fueling mode is
reached at
approximately 800 F. Accordingly, a second fueling mode involving some
distribution of fuel between pilot mixer 102 and main mixer 104 is required.
As
indicated by bar 280, the preferred fueling mode for achieving combustor
temperatures at approximately 800 F is for about 20% of the fuel to be
provided to
pilot mixer 102 and about 80% of the fuel to be provided to main mixer 104.
Utilization of this fueling mode prior to this temperature point (as
represented by
blank portion 282 of bar 280) is possible without adverse outcome, but not
considered
to provide optimum performance of combustor 26. It will also be seen that a
lean
blow out limit 283 for this fueling mode is at a combustor temperature of
approximately 525 F. The second fueling mode is used during a combustor
temperature range of approximately 800 F to approximately 950 F, which is
depicted
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by cross-hatched portion 284 of bar 280. This second fueling mode is then
utilized
during the climb and cruise portions of engine operation.
It is then seen from bar 286 that a third fueling mode is preferred when the
temperature of the combustor inlet air reaches approximately 950 F. It is
preferred
that the third fueling mode preferably include approximately 8% of the fuel
being
provided to pilot mixer 102 and approximately 92% of the fuel being provided
to
main mixer 104. This third temperature stage is represented by cross-hatched
portion
288 of bar 286 and involves a combustor temperature range of approximately 950
F to
approximately 1100 F. Utilization of this third fueling mode prior to this
temperature
point is possible without adverse outcome (see blank portion 290 of bar 286),
but not
considered to provide optimum performance of combustor 26. It will be noted,
however, that a lean blow out limit 292 does exist at approximately 700 F. It
will also
be seen that the second fueling mode (i.e., 20% pilot mixer/80% main mixer)
could be
utilized during this combustor temperature range (approximately 950 F to
approximately 1100 F) without adverse outcome (see blank portion 291 of bar
280),
but it has not been found to provide optimum performance of combustor 26.
Implementation of the third fueling mode is typically done when the greatest
thrust is
required from engine 10, such as during the take-off portion of operation. It
will then
be seen that a fuel pump limit 294 is reached for the third fueling mode
(i.e., 8% pilot
mixer/92% main mixer) at approximately 1100 F.
While there have been described herein what are considered to be preferred and
exemplary embodiments of the present invention, other modifications of these
embodiments falling within the scope of the invention described herein shall
be
apparent to those skilled in the art. For example, it will be understood that
the method
and apparatus of the present invention may be utilized with mixers having
different
configurations. While the mixer shown herein has a pilot mixer with both a
primary
fuel injector and secondary fuel injection ports, it may also be one where
only the
primary fuel injector is provided.
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