U.S. patent application number 11/239376 was filed with the patent office on 2007-04-05 for turbine engine having acoustically tuned fuel nozzle.
This patent application is currently assigned to Solar Turbines Incorporated. Invention is credited to Mario E. Abrau, James W. Blust, Donald James Cramb, Thomas John Chipman Rogers, Christopher Z. Twardochleb.
Application Number | 20070074518 11/239376 |
Document ID | / |
Family ID | 37507653 |
Filed Date | 2007-04-05 |
United States Patent
Application |
20070074518 |
Kind Code |
A1 |
Rogers; Thomas John Chipman ;
et al. |
April 5, 2007 |
Turbine engine having acoustically tuned fuel nozzle
Abstract
A fuel nozzle for a turbine engine having a combustion chamber
is disclosed. The fuel nozzle has a common axis, a body member, and
a barrel member. The fuel nozzle also has a mixing duct and an air
inlet duct, each with predetermined lengths. The fuel nozzle
additionally has a main fuel injection device located between the
air inlet duct and the mixing duct. The main fuel injection device
is configured to introduce a flow of fuel into the barrel member at
a predetermine axial fuel introduction location. The predetermined
axial fuel introduction location and the predetermined length of at
least one of the mixing duct and the air inlet duct are such that a
time-varying fuel to air equivalence ratio at a flame front
downstream of an exit of the mixing duct is less than a
time-averaged fuel to air equivalence ratio when a
naturally-occurring time-varying pressure at the flame front is at
a maximum.
Inventors: |
Rogers; Thomas John Chipman;
(San Diego, CA) ; Twardochleb; Christopher Z.;
(Alpine, CA) ; Blust; James W.; (San Diego,
CA) ; Abrau; Mario E.; (Poway, CA) ; Cramb;
Donald James; (San Diego, CA) |
Correspondence
Address: |
CATERPILLAR/FINNEGAN, HENDERSON, L.L.P.
901 New York Avenue, NW
WASHINGTON
DC
20001-4413
US
|
Assignee: |
Solar Turbines Incorporated
|
Family ID: |
37507653 |
Appl. No.: |
11/239376 |
Filed: |
September 30, 2005 |
Current U.S.
Class: |
60/776 ;
60/737 |
Current CPC
Class: |
F23R 3/286 20130101 |
Class at
Publication: |
060/776 ;
060/737 |
International
Class: |
F23R 3/30 20060101
F23R003/30 |
Claims
1. A fuel nozzle for a turbine engine having a combustion chamber,
comprising: a common axis; a body member disposed about the common
axis; a barrel member located radially outward from the body
member; a mixing duct fluidly communicating the barrel member and
the combustion chamber, and having a predetermined length; an air
inlet duct disposed upstream of the barrel member, having a
predetermined length, and being configured to introduce a flow of
air into the barrel member; and a main fuel injection device
located between the air inlet duct and the mixing duct, the main
fuel injection device configured to introduce a flow of fuel into
the barrel member at a predetermine axial fuel introduction
location, wherein the predetermined axial fuel introduction
location and the predetermined length of at least one of the mixing
duct and the air inlet duct are such that a time-varying fuel to
air equivalence ratio at a flame front downstream of an exit of the
mixing duct is less than a time-averaged fuel to air equivalence
ratio when a naturally-occurring time-varying pressure at the flame
front is at a maximum.
2. The fuel nozzle of claim 1, wherein the predetermined axial fuel
introduction location and the predetermined length of the at least
one of the mixing duct and the air inlet duct are also such that
the time-varying fuel to air equivalence ratio at the flame front
is greater than the time-averaged fuel to air equivalence ratio
when the time-varying pressure at the flame front is at a
minimum.
3. The fuel nozzle of claim 2, wherein the predetermined lengths of
both the mixing duct and air inlet duct are set such that the
time-varying fuel to air equivalence ratio is greater than the
time-averaged fuel to air equivalence ratio when the time-varying
pressure at the flame front is at the minimum and less than the
time-averaged fuel to air equivalence ratio when the time-varying
pressure at the flame front is at the maximum.
4. The fuel nozzle of claim 1, wherein the flow of air introduced
to the barrel member is a time-varying flow and the fuel nozzle
further includes at least one air injection port configured to
inject compressed air into the barrel member at a predetermined
axial location approximately 180 degrees out of phase with the
time-varying flow of air such that attenuation of a pressure wave
traveling from the air inlet duct toward the mixing duct
occurs.
5. The fuel nozzle of claim 4, wherein the at least one air
injection port is a first air injection port and the fuel nozzle
further includes at least a second air injection port axially
aligned with the first air injection port.
6. The fuel nozzle of claim 4, wherein the air inlet duct
introduces a greater amount of air into the fuel nozzle than the at
least one air injection port.
7. The fuel nozzle of claim 4, further including a flow restrictor
located proximal the air inlet duct, the flow restrictor configured
to divert a predetermined portion of the compressed air from the
air inlet duct toward the at least one air injection port.
8. The fuel nozzle of claim 1, wherein the air inlet duct is
substantially straight.
9. The fuel nozzle of claim 1, wherein the mixing duct is
substantially straight.
10. The fuel nozzle of claim 1, wherein the length of the inlet air
duct is such that an axial location of the introduction of the flow
of air is substantially coterminous with the predetermined axial
fuel introduction location.
11. A method of operating a turbine engine, the method comprising:
directing compressed air into the turbine engine via an inlet duct
having a predetermined length; introducing fuel into the turbine
engine at a predetermined axial position downstream of the inlet
duct; mixing the fuel and air within a mixing duct having a
predetermined length; and directing the fuel and air mixture to a
combustion chamber, wherein the predetermined axial fuel
introduction location and the predetermined length of at least one
of the mixing duct and the inlet duct are such that a time-varying
fuel to air equivalence ratio at a flame front downstream of an
exit of the mixing duct is less than a time-averaged fuel to air
equivalence ratio when a naturally-occurring time-varying pressure
at the flame front is at a maximum.
12. The method of claim 11, wherein the predetermined axial fuel
introduction location and the predetermined length of at the least
one of the mixing duct and the inlet duct are also such that the
time-varying fuel to air equivalence ratio at the flame front is
greater than the time-averaged fuel to air equivalence ratio when
the time-varying pressure at the flame front is at a minimum.
13. The method of claim 11, wherein the predetermined lengths of
both the mixing duct and inlet duct are set such that the
time-varying fuel to air equivalence ratio is greater than the
time-averaged fuel to air equivalence ratio when the time-varying
pressure at the flame front is at the minimum and less than the
time-averaged fuel to air equivalence ratio when the time-varying
pressure at the flame front is at the maximum.
14. The method of claim 11, wherein the air directed in to the
turbine engine has a time-varying flow characteristic and the
method further includes injecting compressed air into the turbine
engine at a predetermined axial location approximately 180 degrees
out of phase with the time-varying flow of air such that
attenuation occurs.
15. The method of claim 14, wherein the flow rate of compressed air
through the inlet duct is greater than the flow rate of air
injected at the predetermined axial location within the turbine
engine.
16. The method of claim 11, further including diverting compressed
air from upstream of the inlet duct around the inlet duct to an
injection location downstream of the inlet duct.
17. A turbine engine, comprising: a compressor section configured
to pressurize inlet air; a combustion chamber configured to receive
the pressurized air; and a fuel nozzle configured to direct fuel
into the combustion chamber, the fuel nozzle having: a common axis;
a body member disposed about the common axis; a barrel member
located radially outward from the body member; a mixing duct
fluidly communicating the barrel member and the combustion chamber,
and having a predetermined length; an air inlet duct disposed
upstream of the barrel member, having a predetermined length, and
being configured to introduce a flow of air into the barrel member;
and a main fuel injection device located between the air inlet duct
and the mixing duct, the main fuel injection device configured to
introduce a flow of fuel into the barrel member at a predetermine
axial fuel introduction location, wherein the predetermined axial
fuel introduction location and the predetermined lengths of the
mixing duct and the air inlet duct are such that a time-varying
fuel to air equivalence ratio is greater than a time-averaged fuel
to air equivalence ratio when a time-varying pressure at a flame
front downstream of an exit of the mixing duct is at a minimum and
less than the time-averaged fuel to air equivalence ratio when the
time-varying pressure at the flame front is at a maximum.
18. The turbine engine of claim 17, wherein the flow of air
introduced to the barrel member is a time-varying flow and the fuel
nozzle further includes a plurality of axially aligned air
injection ports configured to inject compressed air into the barrel
member at a predetermined axial location approximately 180 degrees
out of phase with the time-varying flow of air such that
attenuation of a pressure wave traveling from the air inlet duct
toward the mixing duct occurs.
19. The turbine engine of claim 17, wherein the air inlet duct
introduces a greater amount of air into the fuel nozzle than the at
least one air injection port.
20. The turbine engine of claim 17, further including a flow
restrictor located proximal the air inlet duct, the flow restrictor
configured to divert a predetermined portion of the compressed air
from the air inlet duct toward the at least one air injection
port.
21. The turbine engine of claim 17, wherein the air inlet and
mixing ducts are both substantially straight.
Description
TECHNICAL FIELD
[0001] The present disclosure relates generally to a turbine
engine, and more particularly, to a turbine engine having an
acoustically tuned fuel nozzle.
BACKGROUND
[0002] Internal combustion engines, including diesel engines,
gaseous-fueled engines, and other engines known in the art, may
exhaust a complex mixture of air pollutants. These air pollutants
may be composed of gaseous compounds, which may include nitrous
oxides (NOx). Due to increased attention on the environment,
exhaust emission standards have become more stringent and the
amount of NOx emitted to the atmosphere from an engine may be
regulated depending on the type of engine, size of engine, and/or
class of engine.
[0003] It has been established that a well-distributed flame having
a low flame temperature can reduce NOx production to levels
compliant with current emission regulations. One way to generate a
well-distributed flame with a low flame temperature is to premix
fuel and air to a predetermined lean fuel to air equivalence ratio.
However, naturally-occurring pressure fluctuations within the
turbine engine can be amplified during operation of the engine
under these lean conditions. In fact, the amplification can be so
severe that damage and/or failure of the turbine engine can
occur.
[0004] One method that has been implemented by turbine engine
manufacturers to provide lean fuel/air operational conditions
within a turbine engine while minimizing the harmful vibrations
generally associated with lean operation is described in U.S. Pat.
No. 6,698,206 (the '206 patent) issued to Scarinci et al. on Mar.
2, 2004. The '206 patent describes a turbine engine having a
primary combustion zone, a secondary combustion zone, and a
tertiary combustion zone. Each of the combustion zones is supplied
with premixed fuel and air by respective mixing ducts and a
plurality of axially spaced-apart air injection apertures. These
apertures reduce the magnitude of fluctuations in the lean fuel to
air equivalence ratio of the fuel and air mixtures supplied into
the mixing zones, thereby reducing the harmful vibrations.
[0005] Although the method described in the '206 patent may reduce
some harmful vibrations associated with a low NOx-emitting turbine
engine, it may be expensive and insufficient. In particular, the
many apertures associated with each of the combustion zones
described in the '206 patent may drive up the cost of the turbine
engine. In addition, because the reduction of vibration within the
turbine engine of the '206 patent does not rely upon strategic
placement of the apertures according to acoustic tuning specific to
the particular turbine engine, the reduction of vibration may be
limited and, in some situations, insufficient.
[0006] The disclosed fuel nozzle is directed to overcoming one or
more of the problems set forth above.
SUMMARY OF THE INVENTION
[0007] In one aspect, the present disclosure is directed to a fuel
nozzle for a turbine engine having a combustion chamber. The fuel
nozzle includes a common axis, a body member disposed about the
common axis, and a barrel member located radially outward from the
body member. The fuel nozzle also includes a mixing duct fluidly
communicating the barrel member and the combustion chamber, and an
air inlet duct disposed upstream of the barrel member. The air
inlet duct is configured to introduce a flow of air into the barrel
member. Each of the air inlet duct and mixing duct have
predetermined lengths. The fuel nozzle further includes a main fuel
injection device located between the air inlet duct and the mixing
duct. The main fuel injection device is configured to introduce a
flow of fuel into the barrel member at a predetermined axial fuel
introduction location. The predetermined axial fuel introduction
location and the predetermined length of at least one of the mixing
duct and the air inlet duct are such that a time-varying fuel to
air equivalence ratio at a flame front downstream of the mixing
duct is less than a time-averaged fuel to air equivalence ratio
when a naturally-occurring time-varying pressure at the flame front
is at a maximum.
[0008] In another aspect, the present disclosure is directed to a
method of operating a turbine engine. The method includes directing
compressed air into the turbine engine via an inlet duct having a
predetermined length. The method also includes introducing fuel
into the turbine engine at a predetermined axial position
downstream of the inlet duct, and mixing the fuel and air within a
mixing duct having a predetermined length. The method further
includes directing the fuel and air mixture to a combustion
chamber. The predetermined axial fuel introduction location and the
predetermined length of at least one of the mixing duct and the
inlet duct are such that a time-varying fuel to air equivalence
ratio at a flame front downstream of an exit of the mixing duct is
less than a time-averaged fuel to air equivalence ratio when a
naturally-occurring time-varying pressure at the flame front is at
a maximum.
BRIEF DESCRIPTION OF THE DRAWINGS
[0009] FIG. 1 is a cutaway-view illustration of an exemplary
disclosed turbine engine;
[0010] FIG. 2 is a cross-sectional illustration of an exemplary
disclosed fuel nozzle for the turbine engine of FIG. 1; and
[0011] FIG. 3 is a pictorial representation of an exemplary
disclosed operation of the fuel nozzle of FIG. 2.
DETAILED DESCRIPTION
[0012] FIG. 1 illustrates an exemplary turbine engine 10. Turbine
engine 10 may be associated with a stationary or mobile work
machine configured to accomplish a predetermined task. For example,
turbine engine 10 may embody the primary power source of a
generator set that produces an electrical power output or of a
pumping mechanism that performs a fluid pumping operation. Turbine
engine 10 may alternatively embody the prime mover of an
earth-moving machine, a passenger vehicle, a marine vessel, or any
other mobile machine known in the art. Turbine engine 10 may
include a compressor section 12, a combustor section 14, a turbine
section 16, and an exhaust section 18.
[0013] Compressor section 12 may include components rotatable to
compress inlet air. Specifically, compressor section 1 2 may
include a series of rotatable compressor blades 22 fixedly
connected about a central shaft 24. As central shaft 24 is rotated,
compressor blades 22 may draw air into turbine engine 10 and
pressurize the air. This pressurized air may then be directed
toward combustor section 14 for mixture with a liquid and/or
gaseous fuel. It is contemplated that compressor section 12 may
further include compressor blades (not shown) that are separate
from central shaft 24 and remain stationary during operation of
turbine engine 10.
[0014] Combustor section 14 may mix fuel with the compressed air
from compressor section 12 and combust the mixture to create a
mechanical work output. Specifically, combustor section 14 may
include a plurality of fuel nozzles 26 annularly arranged about
central shaft 24, and an annular combustion chamber 28 associated
with fuel nozzles 26. Each fuel nozzle 26 may inject one or both of
liquid and gaseous fuel into the flow of compressed air from
compressor section 12 for ignition within combustion chamber 28. As
the fuel/air mixture combusts, the heated molecules may expand and
move at high speed into turbine section 16.
[0015] As illustrated in the cross-section of FIG. 2, each fuel
nozzle 26 may include components that cooperate to inject gaseous
and liquid fuel into combustion chamber 28. Specifically, each fuel
nozzle 26 may include a barrel housing 34 connected on one end to
an air inlet duct 35 for receiving compressed air, and on the
opposing end to a mixing duct 37 for communication of the fuel/air
mixture with combustion chamber 28. Fuel nozzle 26 may also include
a central body 36, a pilot fuel injector 38, and a swirler 40.
Central body 36 may be disposed radially inward of barrel housing
34 and aligned along a common axis 42. Pilot fuel injector 38 may
be located within central body 36 and configured to inject a pilot
stream of pressurized fuel through a tip end 44 of central body 36
into combustion chamber 28 to facilitate engine starting, idling,
cold operation, and/or lean burn operations of turbine engine 10.
Swirler 40 may be annularly disposed between barrel housing 34 and
central body 36.
[0016] Barrel housing 34 may embody a tubular member having a
plurality of air jets 46. Air jets 46 may be co-aligned at a
predetermined axial position along the length of barrel housing 34.
This predetermined axial position may be set during manufacture of
turbine engine 10 to attenuate a time-varying flow of air entering
fuel nozzle 26 via air inlet duct 35. It is contemplated that air
jets 46 may be located at any axial position along the length of
barrel housing 34 and may vary from engine to engine or from one
class or size of engine to another class or size of engine
according to attenuation requirements. Air jets 46 may receive
compressed air from compressor section 12 by way of one or more
fluid passageways (not shown) external to barrel housing 34.
[0017] Air inlet duct 35 may embody a tubular member configured to
axially direct compressed air from compressor section 12 (referring
to FIG. 1) to barrel housing 34, and to divert a portion of the
compressed air to air jets 46. Specifically, air inlet duct 35 may
include a central opening 48 and a flow restrictor 50 located
within central opening 48 at an end opposite barrel housing 34. In
one example, flow restrictor 50 may embody a blocker ring extending
inward from the interior surface of air inlet duct 35. The radial
distance that flow restrictor 50 protrudes into central opening 48
may determine the amount of compressed air diverted around air
inlet duct 35 to air jets 46 during operation of turbine engine 10.
The amount of air diverted to air jets 46 may be less than the
amount of air passing through air inlet duct 35. The geometry of
air inlet duct 35 may such that pressure fluctuations within fuel
nozzle 26 may be minimized to provide for piece-wise uniform flow
through air inlet duct 35. In one example, air inlet duct 35 may be
generally straight and may have a predetermined length. The
predetermined length of air inlet duct 35 may be set during
manufacture of turbine engine 10 according to an axial fuel
introduction location and a naturally-occurring pressure
fluctuation with combustion chamber 28. The method of determining
and setting the length of air inlet duct 35 will be discussed in
more detail below.
[0018] Mixing duct 37 may embody a tubular member configured to
axially direct the fuel/air mixture from fuel nozzle 26 into
combustion chamber 28. In particular, mixing duct 37 may include a
central opening 52 that fluidly communicates barrel housing 34 with
combustion chamber 28. The geometry of mixing duct 37 may be such
that pressure fluctuations within fuel nozzle 26 are minimized to
provide for piece-wise uniform flow through air inlet duct 35. In
one example, mixing duct 37 may be generally straight and may have
a predetermined length. Similar to air inlet duct 35, the
predetermined length of mixing duct 37 may be set during
manufacture of turbine engine 10 according to an axial fuel
introduction location and the naturally-occurring pressure
fluctuation within combustion chamber 28. The method of determining
and setting the length of mixing duct 37 will be discussed in more
detail below.
[0019] Swirler 40 may be situated to radially redirect an axial
flow of compressed air from air inlet duct 35. In particular,
swirler 40 may embody an annulus having a plurality of connected
vanes 54 located within an axial flow path of the compressed air.
As the compressed air contacts vanes 54, it may be diverted in a
radially inward direction. It is contemplated that vanes 54 may
extend from barrel housing 34 radially inward directly toward
common axis 42 or, alternatively, to a point cantered off-center
from common axis 42. It is also contemplated that vanes 54 may be
straight or twisted along a length direction and tilted at an angle
relative to an axial direction of common axis 42.
[0020] Vanes 54 may facilitate fuel injection within barrel housing
34. In particular, some or all of vanes 54 may each include a
liquid fuel jet 56 and a plurality of gaseous fuel jets 58. It is
contemplated that any number or configuration of vanes 54 may
include liquid fuel jets 56. The location of vanes 54 along common
axis 42 and the resulting axial fuel introduction point within fuel
nozzle 26 may vary and be set to, in combination with specific
time-varying air flow characteristics, attenuate the
naturally-occurring pressure fluctuation within combustion chamber
28. The method of determining and setting the axial fuel
introduction point will be discussed in more detail below.
[0021] Gaseous fuel jets 58 may provide a substantially constant
mass flow of gaseous fuel such as, for example, natural gas,
landfill gas, bio-gas, or any other suitable gaseous fuel to
combustion chamber 28. In particular, gaseous fuel jets 58 may
embody restrictive orifices situated along a leading edge of each
vane 54. Each of gaseous fuel jets 58 may be in communication with
a central fuel passageway 59 within the associated vane 54 to
receive gaseous fuel from an external source (not shown). The
restriction at gaseous fuel jets 58 may be the greatest restriction
applied to the flow of gaseous fuel within fuel nozzle 26, such
that a substantially continuous mass flow of gaseous fuel from
gaseous fuel jets 58 may be ensured.
[0022] Combustion chamber 28 (referring to FIG. 1) may house the
combustion process. In particular, combustion chamber 28 may be in
fluid communication with each fuel nozzle 26 and may be configured
to receive a substantially homogenous mixture of fuel and
compressed air. The fuel/air mixture may be ignited and may fully
combust within combustion chamber 28. As the fuel/air mixture
combusts, hot expanding gases may exit combustion chamber 28 and
enter turbine section 16.
[0023] Turbine section 16 may include components rotatable in
response to the flow of expanding exhaust oases from combustor
section 14. In particular, turbine section 16 may include a series
of rotatable turbine rotor blades 30 fixedly connected to central
shaft 24. As turbine rotor blades 30 are bombarded with high-energy
molecules from combustor section 14, the expanding molecules may
cause central shaft 24 to rotate, thereby converting combustion
energy into useful rotational power. This rotational power may then
be drawn from turbine engine 10 and used for a variety of purposes.
In addition to powering various external devices, the rotation of
turbine rotor blades 30 and central shaft 24 may drive the rotation
of compressor blades 22.
[0024] Exhaust section 18 may direct the spent exhaust from
combustor and turbine sections 14, 16 to the atmosphere. It is
contemplated that exhaust section 18 may include one or more
treatment devices configured to remove pollutants from the exhaust
and/or attenuation devices configured to reduce the noise
associated with turbine engine 10, if desired.
[0025] FIG. 3 illustrates an exemplary relationship between the
length of air inlet duct 35, the length of mixing duct 37, the
axial fuel introduction point within barrel housing 34 resulting
from the position of swirler 40 along common axis 42, and the
naturally-occurring pressure fluctuation stemming from a flame
front 67 within combustion chamber 28. FIG. 3 will be discussed in
more detail below.
INDUSTRIAL APPLICABILITY
[0026] The disclosed fuel nozzle may be applicable to any turbine
engine where reduced vibrations within the turbine engine are
desired. Although particularly useful for low NOx-emitting engines,
the disclosed fuel nozzle may be applicable to any turbine engine
regardless of the emission output of the engine. The disclosed fuel
nozzle may reduce vibrations by acoustically attenuating a
naturally-occuring pressure fluctuation within a combustion chamber
of the turbine engine. The operation of fuel nozzle 26 will now be
explained.
[0027] During operation of turbine engine 10, air may be drawn into
turbine engine 10 and compressed via compressor section 12
(referring to FIG. 1). This compressed air may then be axially
directed into combustor section 14 and against vanes 54 of swirler
40, where the flow may be redirected radially inward. As the flow
of compressed air is turned to flow radially inward, liquid fuel
may be injected from liquid fuel jets 56 for mixing prior to
combustion. Alternatively or additionally, gaseous fuel may be
injected from gaseous fuel jets 58 for mixing with the compressed
air prior to combustion. As the mixture of fuel and air enters
combustion chamber 28, it may ignite and fully combust. The hot
expanding exhaust gases may then be expelled into turbine section
16, where the molecular energy of the combustion gases may be
converted to rotational energy of turbine rotor blades 30 and
central shaft 24.
[0028] FIG. 3 illustrates the time-varying flow characteristics of
fuel and air entering fuel nozzle 26 and their effects on the
naturally-occuring pressure fluctuations within combustion chamber
28. In particular, FIG. 3 illustrates a first curve 60, a second
curve 62, a third curve 64, and a plurality of pressure pulses 66.
First curve 60 may represent the time-varying flow of compressed
air entering fuel nozzle 26 via air inlet duct 35. Second curve 62
may represent the time-varying flow of fuel flow entering fuel
nozzle 26 via liquid and/or gaseous fuel jets 56, 58. Third curve
64 may represent the time-varying fuel to air equivalence ratio
.PHI. (e.g., the instantaneous ratio of the amount of fuel within
any axial plane along the length of fuel nozzle 26 to the amount of
air in the same axial plane). Pressure pulses 66 may represent a
wave of pressure traveling from combustion chamber 28 in a reverse
direction toward air inlet duct 35 as a result of combustion within
combustion chamber 28.
[0029] Pressure pulses 66 may affect the time-varying
characteristic of first, second, and third curves 60-64.
Specifically, as pressure pulses 66 travel in the reverse direction
within fuel nozzle 26 and reach liquid and gaseous fuel injectors
56, 58 and the entrance to air inlet duct 35, the pressure of each
pulse may cause the flow rate of fuel and air entering fuel nozzle
26 to vary. These varying flow rates correspond to the amplitude
variations of first and second curves 60, 62 illustrated in FIG. 3,
which equate to the varying amplitude and phase angle of third
curve 64. When the value of .PHI. at the point of combustion within
combustion chamber 28 is high compared to a time average value of
.PHI., the heat release and resulting pressure wave within
combustion chamber 28 may be high. Likewise, when the value of
.PHI. at the point of combustion within combustion chamber 28 is
low compared to the time average value of .PHI., the heat release
and resulting pressure wave within combustion chamber 28 may be
low.
[0030] Damage may occur when the phase angle of third curve 64 and
the wave of pressure pulses 66 near alignment. That is, when the
value of .PHI. entering combustion chamber 28 is high compared to
the time average of .PHI. and enters combustion chamber 28 at about
the same time that a pressure pulse 66 initiates from a flame front
with combustion chamber 28, resonance may be attained. Likewise, if
the value of .PHI. entering combustion chamber 28 is low compared
to the time average of .PHI. and enters combustion chamber 28 at a
time between the intiation of pressure pulses 66, resonance may be
attained. It may be possible that this resonance could amplify
pressure pulses 66 to a damaging magnitude.
[0031] Damage may be prevented when third curve 64 and the wave of
pressure pulses 66 are out of phase. In particular, if the value of
.PHI. entering combustion chamber 28 is low compared to the time
average of .PHI. and enters combustion chamber 28 at the same time
that a pressure pulse 66 initiates from a flame front within
combustion chamber 28, attenuation of pressure pulse 66 may be
attained. Likewise, if the value of .PHI. entering combustion
chamber 28 is high compared to the time average of .PHI. and enters
combustion chamber 28 at a time between the imitation of pressure
pulses 66, attenuation may be attained. Attenuation could lower the
magnitude of pressure pulses 66, thereby minimizing the likelihood
of damage to turbine engine 10.
[0032] The phase angle and magnitude of .PHI. may be affected by
the length of air inlet duct 35, the length of mixing duct 37, the
axial fuel introduction point, and the axial location of air jets
46. Specifically, by increasing the length of air inlet duct 35
(e.g., extending the entrance of air inlet duct 35 leftward, when
viewed in FIG. 2), the phase angle of first curve 60 may likewise
shift to the left. In contrast, by decreasing the length of air
inlet duct 35 (e.g., moving the entrance of air inlet duct 35 to
the right, when viewed in FIG. 2), the phase angle of first curve
60 may likewise move to the right. In fact, if the length of air
inlet duct 35 becomes so short that the introduction of air is
substantially coterminous with the introduction of fuel via gaseous
fuel jets 58 and the pressure drops across flow restrictor 50 and
gaseous fuel jets 58 are substantially constant, the phase angle
and amplitude differences between first and second curves 60, 62
may be nearly zero, resulting in a substantially constant value of
.PHI.. In addition, by extending the length of mixing duct 37
(e.g., extending the exit of mixing duct 37 rightward, when viewed
FIG. 2), the phase angle of first curve 60 may move to the left. By
decreasing the length of mixing duct 37 (e.g., moving the exit of
mixing duct 37 leftward, when viewed in FIG. 2), the phase angle of
first curve 60 may move to the right. By moving the location of
swirler 40 left or right and, in doing so, the axial introduction
point of gaseous and liquid fuel left or right, the phase angle of
second curve 62 may mimic the same shifts. As the phase angle of
one or both of first and second curves 60, 62 shifts, the phase
angle and amplitude of third curve 64 may be affected. In this
manner, the value of .PHI. entering combustion chamber 28 can be
acoustically tuned to attenuate the naturally-occuring pressure
pulses 66 of a specific engine or specific class or size of engine.
It is contemplated that only one or both of the lengths of air
inlet duct 35 and mixing duct 37 may be modified to attenuate the
naturally-occurring pressure pulses 66.
[0033] Further reduction in the magnitude of pressure pulses 66 may
be attained by providing a substantially time-constant value of
.PHI.. One way to reduce the variation in the value of .PHI. may be
to reduce the time-varying characteristic of first and/or second
curves 60, 62. The time-varying characteristic of gaseous fuel
introduced into combustion chamber 28 via gaseous fuel jets 58 may
be reduced by way of the restriction at the surface of gaseous fuel
jets 58. This restriction may increase the pressure drop across
gaseous fuel jets 58 to a magnitude at which the pressure
fluctuations within fuel nozzle 26 may have little affect on the
flow of fuel through gaseous fuel jets 58. Another way to reduce
the vibrations may be realized through the use of air jets 46. In
particular, as seen in FIG. 3, when pulses of compressed air are
introduced at a specific location within fuel nozzle 26 and at a
timing out of phase with first curve 60, the time-varying
characteristic of air entering combustion chamber 28 may be
attenuated. In one example, the pulses of compressed air may be
injected by air jets 46 substantially 180 degrees out of phase with
first curve 60. The affect of the injected pulses of air can be
seen in FIG. 3; as the flow of compressed air entering barrel
housing 34 via air inlet duct 35 passes in proximity to air jets
46, the amplitude of first curve 60 may be reduced.
[0034] Several advantages over the prior art may be associated with
fuel nozzle 26 of turbine engine 10. Specifically, because the
length of air inlet duct 35, the length of mixing duct 37, and the
axial fuel introduction point of turbine engine 10 may be selected
specifically to attenuate the naturally-occurring pressure pulses
of combustion chamber 28, harmful vibrations of turbine engine 10
may be greatly reduced. This acoustic tuning of turbine engine 10
may be more successful at reducing vibration than the random
placement of apertures in an attempt to create non-resonating
turbulence. In addition, these reductions in vibration may be
attained with minimal changes to existing hardware, resulting in
lower component costs of turbine engine 10.
[0035] It will be apparent to those skilled in the art that various
modifications and variations can be made to the disclosed fuel
nozzle. Other embodiments will be apparent to those skilled in the
art from consideration of the specification and practice of the
disclosed fuel nozzle. It is intended that the specification and
examples be considered as exemplary only, with a true scope being
indicated by the following claims and their equivalents.
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