U.S. patent number 3,851,462 [Application Number 05/375,248] was granted by the patent office on 1974-12-03 for method for reducing turbine inlet guide vane temperatures.
This patent grant is currently assigned to United Aircraft Corporation. Invention is credited to Alexander Vranos.
United States Patent |
3,851,462 |
Vranos |
December 3, 1974 |
METHOD FOR REDUCING TURBINE INLET GUIDE VANE TEMPERATURES
Abstract
The method for reducing turbine inlet guide vane temperatures in
a gas turbine engine includes the step of oscillating the fuel
delivery into the combustion chamber. In a preferred embodiment the
fuel spray cone angle is varied at a high frequency equivalent to
the frequency of one of the harmonics or subharmonics of the
natural acoustic pressure oscillations of the combustion chamber.
Apparatus for this method is also provided.
Inventors: |
Vranos; Alexander (Rockville,
CT) |
Assignee: |
United Aircraft Corporation
(East Hartford, CT)
|
Family
ID: |
23480124 |
Appl.
No.: |
05/375,248 |
Filed: |
June 29, 1973 |
Current U.S.
Class: |
60/776; 60/39.37;
60/742; 60/39.281; 60/740 |
Current CPC
Class: |
F23R
3/06 (20130101); F23R 3/28 (20130101); Y02T
50/675 (20130101); Y02T 50/60 (20130101) |
Current International
Class: |
F23R
3/28 (20060101); F23R 3/06 (20060101); F23R
3/04 (20060101); F02c 007/12 (); F02c 007/22 () |
Field of
Search: |
;60/39.28R,39.36,39.9R,39.82R,39.46,39.37,39.74R,39.74B,39.06,39.65
;431/12 |
References Cited
[Referenced By]
U.S. Patent Documents
Primary Examiner: Croyle; Carlton R.
Assistant Examiner: Olsen; Warren
Attorney, Agent or Firm: Revis; Stephen E.
Claims
Having thus described typical embodiments of my invention, that
which I claim as new and desire to secure by Letters Patent of the
United States is:
1. In a gas turbine engine including a combustion chamber and a
plurality of circumferentially spaced turbine inlet guide vanes
positioned at the downstream end of the combustion chamber, said
combustion chamber including at least one burner can having at
least one row of circumferentially spaced combustion air holes
through one wall thereof, the method of reducing the maximum
temperature reached by said vanes including the steps of:
spraying fuel into the burner can; and
oscillating the fuel delivery into the burner can with a period of
oscillation shorter than the turbine inlet guide vane thermal
response time and at a selected frequency which is the frequency of
one of the harmonics or subharmonics of the natural acoustic
pressure oscillations of the combustion chamber.
2. The method for reducing maximum vane temperature according to
claim 1 wherein the step of oscillating the fuel delivery includes
oscillating the spray pattern of the fuel with said period of
oscillation and at said selected frequency as it is delivered into
the burner can.
3. The method for reducing maximum vane temperature according to
claim 2 wherein the step of spraying fuel into the burner can
includes spraying fuel into the burner can in the form of a cone,
and the step of oscillating the spray pattern includes varying the
cone angle of the spray.
4. The method of reducing maximum vane temperature according to
claim 1 wherein the step of oscillating the fuel delivery includes
oscillating the amount of fuel delivered into the burner can
between larger and smaller amounts at said selected frequency.
Description
BACKGROUND OF THE INVENTION
1. Field of the Invention
This invention relates to a method and apparatus for reducing
turbine inlet guide vane temperatures in a gas turbine engine.
2. Description of the Prior Art
Advancement in gas turbine engine design is often hampered by
materials limitations in the turbine section which is subjected to
high combustor exit temperatures. Materials development is
sometimes hard pressed to keep pace with rising temperatures. The
turbine inlet guide vane is of particular concern since it must be
able to withstand the brunt of extremely high temperatures exiting
from the combustion chamber. Prior art solutions to these materials
problems have been focused in two major areas: One is to provide
elaborate cooling schemes for the vanes and the other is to reduce
the gas temperatures exiting from the combustion chamber. The
former solution is expensive to apply and is usually used as a last
resort. The latter solution is really not a solution at all since
temperatures must increase as engines become larger and
requirements become more demanding.
SUMMARY OF THE INVENTION
Accordingly, an object of the present invention is to reduce
turbine inlet guide vane temperatures.
Another object of the present invention is to reduce turbine inlet
guide vane temperatures without reducing combustion chamber
temperatures.
A further object of the present invention is to reduce turbine
inlet guide vane temperatures without flowing cooling fluid over or
within the turbine inlet guide vanes.
In one form the invention is the method for reducing the surface
temperatures of a turbine inlet guide vane and includes the step of
oscillating the fuel delivery into the combustion chamber and
coupling the oscillations with the natural acoustic pressure
oscillations of the combustion chamber. The fuel delivery can be
oscillated in several ways, such as by feeding the fuel into the
front end of the combustion chamber in alternately large and small
amounts at the selected frequency. Another technique is by
oscillating the fuel spray pattern at the selected frequency such
as by varying the fuel spray cone angle.
In the latter method, varying the cone spray angle will vary the
concentration of fuel within the recirculation zone from a high to
a lower value, thereby resulting in a heat release rate which will
vary from a high to a lower value. The frequency of oscillation in
the cone spray angle may be set so that the heat release rate is
coupled with combustion chamber acoustics (i.e., the frequency is
the same as one of the harmonics or subharmonics of the natural
acoustic pressure oscillation within the combustion chamber) in a
manner to be hereinafter explained in more detail in the
description of the preferred embodiments. By coupling the heat
release rate with the burner acoustics, cool air jets entering
through the burner liner can be made to oscillate at a high rate of
speed across the surfaces of the turbine inlet guide vanes thereby
preventing a hot streak of gases from being concentrated on a
single small area of a guide vane for a period of time long enough
to do damage. The sweeping action of the air jets results in a
reduced overall vane average temperature.
In a preferred embodiment of the present invention a fuel spray
nozzle having inner and outer annular passages with a common source
and feedback loops between the passages causes the fuel to be
driven automatically through first one passage and then the other
passage in an alternating fashion, each passage spraying the fuel
into the combustion chamber at a different spray angle. The
preferred feedback technique used to oscillate the fuel between the
inner and outer conical passages is shown in the Bodine U.S. Pat.
No. 3,111,931; however, it has not been used in a gas turbine
engine fuel nozzle or for the purposes of the present
invention.
The foregoing and other objects, features and advantages of the
present invention will become more apparent in the light of the
following detailed description of the preferred embodiments thereof
as illustrated in the accompanying drawing.
BRIEF DESCRIPTION OF THE DRAWING
FIG. 1 is a cross sectional view, partly schematic, showing a
combustion chamber incorporating the features of the present
invention.
FIG. 2 is a cross sectional view, partly schematic, of the burner
can shown in FIG. 1.
FIG. 3 is an enlarged, illustrative cross sectional view of the
downstream end of the burner can shown in FIGS. 1 and 2.
FIG. 4 is a cross sectional view, partly schematic, of the upstream
portion of a burner can according to one embodiment of the present
invention.
FIG. 5 is a cross sectional view taken along the line 5--5 of FIG.
4.
FIG. 6 is a cross sectional view of an alternate embodiment of the
nozzle shown in FIGS. 4 and 5.
FIG. 7 is an offset cross sectional view taken along the line 7--7
of FIG. 6.
FIG. 8 is a cross sectional view, partly schematic, of the upstream
portion of a burner can incorporating another embodiment of the
present invention.
FIG. 9 is an illustrative cross sectional view, partly schematic,
of an annular combustion chamber assembly according to the present
invention.
FIG. 10 is a cross sectional view taken along the line 10--10 of
FIG. 9.
DESCRIPTIONS OF THE PREFERRED EMBODIMENTS
Before entering into a detailed discussion of the methods and
apparatus of the present invention, it is important to understand
some of the more basic mechanisms of the gas turbine engine
combustion process which affect or might have an effect on the
temperatures reached by the turbine inlet guide vanes at the exit
of the combustion chamber. Although in the foregoing sentence the
term "basic mechanisms" is used, it should be made clear at this
time that "basic" does not necessarily mean "know to those skilled
in the art;" on the contrary, the following discussion is based on
relatively new theories which have been supported by experimental
testing and by analysis of current gas turbine engine combustion
chambers. It is this new insight into the combustion process that
has lead to the development of the present invention; and it is for
that reason that at least a brief description of these theories are
necessary for a complete understanding of the present
invention.
Consider the gas turbine engine combustion chamber of the can-type
shown in FIG. 1. The combustion chamber 10 comprises an annular
duct 12 formed by an outer casing 14 and an inner casing 16. The
forward end of the annular duct 12 includes a diffuser section 18.
Circumferentially spaced within the duct 12 are a plurality of
substantially cylindrical burner cans 20. Each burner can 20 has a
dome-shaped upstream end 22 and an open downstream end 24 which
interconnects with a transition duct 26 forming an annular exit
zone 28. Immediately downstream of the annular exit zone 28 are
positioned a plurality of turbine inlet guide vanes 30. Centrally
located within each dome portion 22 is a fuel nozzle 32 for
spraying fuel into the upstream portion 34 of the can 20.
Surrounding each fuel nozzle 32 are swirler vanes 36 which admit
air into the front end portion 34 to mix with the fuel. Each burner
can 20 also includes several rows of holes through its wall for
admitting additional air into the can 20 at various locations along
its length. Although there may be several rows of holes, for
purposes of the present discussion, it is sufficient that there be
a row of primary combustion air holes 38 for supplying air into the
upstream portion 34 or primary combustion zone of the burner can
20, and a downstream row of holes 40 for supplying dilution and
cooling air to the downstream portion 42 or cooling and mixing zone
of the burner can 20.
As is known to those skilled in the art, when a fluid is forced to
flow through a tube or a duct, the duct will display certain
resonant acoustic pressure modes of various determinable
frequencies which will depend upon whether or not the duct is open
or closed, the length and shape of the duct, and the velocity of
sound in the media flowing through the duct. In the combustion
chamber 10 of FIG. 1 the annular duct 12 and each burner can 20 act
as partially closed tubes. Their longitudinal mode frequencies can
be calculated and their resonances will be stable. For the purposes
of this discussion, it is only necessary to consider the
longitudinal modes since they are the most probable modes because
the radial and tangential modes lack an efficient method of
excitation due to the geometries and boundary conditions of the
combustion chamber. It should also be noted, that because the duct
12 and the cans 20 are coupled to each other by means of holes 38
and 40, that there will exist in these ducts frequencies of the
various acoustic modes which are the difference between the
frequencies in the duct 12 and in the cans 20. From the foregoing
it is apparent that the pressure drop across the wall of the can 20
at any particular axial location, such as for example, at the
primary combustion air holes 38, varies with time as the pressure
waves move through the duct 12 and the burner cans 20. The pressure
drop will fluctuate more when the pressure waves are out of phase
than when they are in phase. FIG. 2 is a simplified drawing of the
burner can of FIG. 1 and is illustrative of the effect that this
varying pressure drop has upon the air entering the combustion
chamber through the primary combustion air holes 38. When the
pressure drop is large, the air jet has a higher velocity and
penetrates well into the burner can 20 as indicated by the arrow
44. On the other hand, when the pressure drop is less the air jet
is more easily deflected by the gases within the burner can 20
traveling downstream and so it does not penetrate as far into the
combustion chamber 20. The arrow 46 is representative of this
condition. Thus, the air jet oscillates between the positions 44
and 46 at a frequency and in a manner determined in part by the
pressure pulses within the duct 12 and the can 20 and whether or
not these pulses are in or out of phase.
The same phenomenon occurs at the dilution air holes 40 as
represented by the arrows 48 and 50.
FIG. 3 is an enlarged view of the downstream end of FIG. 2 showing
the dilution air holes 40, the transition duct 26 and the turbine
inlet guide vanes 30. The arrows 48, 50, representing the two
positions of an air jet entering through the holes 40 are more
realistically shown as bands of air 48', 50' in FIG. 3. The
position 50' is shown in dotted outline while the position 48' is
shown in phantom. The air jet is generally circular in cross
section and is cooler in temperature near the center than around
its periphery. Thus, as the jet sweeps back and forth from the
position 48' to the position 50', a single point on the surface of
the vane 30 is subject alternately to hotter and then cooler air.
It should be evident that this is desirable, for if the jet did not
oscillate then the portion of the vane 30 impinged upon by the
hotter outer periphery of the air jet would reach a temperature
substantially the same as the temperature of that portion of the
air jet. Oscillation of the air jet results in a reduced average
temperature. It is believed that this phenomenon occurs to a
greater or lesser degree in current gas turbine engines, but no
attempts have been made to knowingly take advantage of it in the
past, probably because it has not been understood.
A further consideration in this regard is that the vane material,
and for that matter any material, does not respond instantaneously
to a change in temperature of the surrounding fluid. In other
words, the vane will have what is known as a thermal response time
which is related to the speed at which the temperature of the vane
changes at a particular point, given a sudden change in temperature
of the surrounding fluid. The vane thermal response time is herein
defined as the time it takes a vane to reach the temperature of the
surrounding fluid when suddenly exposed to that fluid. The thermal
response time of a vane will depend on several factors including
the thickness of the vane in the area of concern, the composition
of the base material of the vane, the composition of any coating on
the vane, and the magnitude of the temperature change. Because of
this thermal response time the speed with which the air jets
traverse the surface of the vane will have an effect on the maximum
temperature that the vane reaches. If the oscillations of the air
jets are fast enough then the vane will not be able to reach the
temperature of the hottest portion of the air jet. Although in
FIGS. 2 and 3 only the air jets from the dilution holes 40 are
shown impinging upon the vane 30, the same phenomenon occurs with
the air jets entering the primary combustion air holes 38.
From the foregoing it becomes apparent that it is desirable to have
the air jets oscillate at a high frequency in a manner best suited
to maintain the vane average temperature as low as possible. Not
any type of oscillation of the air jets will be effective. As a
matter of fact, most every jet engine will have oscillations of
their air jets; however, it is important to recognize that
generally these oscillations will be random depending on the
frequencies of the pressure pulses occurring inside and outside of
the burner can. There may be relatively long periods of time when
the pressure pulses are in such a phase that the pressure drop
across the liner remains relatively constant. If this occurs for
more than several milliseconds, the air jet may remain stationary
upon the turbine inlet guide vane 30 and cause damage due to
excessive concentrations of heat over a long period of time. It is
thus desirable to control the frequencies of the pressure pulses
inside and outside the duct and to use other methods to cause
oscillations at a high, known frequency. The present invention is
concerned with several methods and apparatus for doing just
that.
It should also be mentioned that the depth of penetration of the
air jets represented by the arrows 44, 46 is also dependent upon
the velocity of the burning gases within the burner can 20. The
velocity of the gases in the burner can 20 are in turn dependent
upon the heat release rate of the combustion process. A high heat
release rate is followed by high velocities and a low heat release
rate is followed by lower velocities. Thus, if a proper cycle of
high and low heat release rates can be established that will couple
with the acoustics of the duct 12 and the cans 20, then the
oscillatory motion of the air jets can be reinforced or amplified
and can be controlled.
In can type burners such as the burner can 12 the majority of
combustion takes place in the primary combustion zone 34. In the
past it has been thought that combustion of the gases within the
primary combustion zone occurs in a continuous, steady state
fashion, it is now felt that combustion occurs as a series of high
frequency explosions. The concentration of fuel within the primary
combustion zone builds up until an explosion occurs resulting in a
rapid increase in pressure in the primary combustion zone 34
followed by a rapid increase in the velocity of gases moving
downstream which tends to move the air jets to the position of the
arrow 46. Immediately after the explosion there is a sudden
decrease in the pressure in the primary combustion zone resulting
in the air jets moving back toward the primary combustion zone to
the position of the arrow 44.
This invention suggests modifying the delivery of fuel into the
primary combustion zone so that alternately there is a high and
then a low concentration of fuel within the primary combustion zone
34. For example alternately increasing and then decreasing the
amount of fuel supplied to the combustion zone causes the heat
release rate and thus the pressure within the primary combustion
zone 34 to alternately increase and decrease. This increase and
decrease in the pressure caused by oscillation of the fuel supply
intensifies the swings in pressure occurring as a result of the
normal combustion process. In this manner the motion of the air
jets may be amplified and the frequency of their oscillations
controlled. This same effect will be felt at the row of holes 40
and at other holes located elsewhere along the burner can.
In the embodiment of this invention shown in FIG. 1, fuel is
supplied to the nozzles 32 from a source represented by the box 33.
Prior to entering the nozzles 32 the fuel passes through
oscillation means represented by the box 35. The oscillation means
35 may operate in several ways. For example, it may simply
alternately increase and then decrease the fuel supply to the
nozzles 32 at a frequency which will result in oscillations of the
heat release rate that are coupled with the natural acoustic
frequencies of the pressure pulses within the duct 12 and the can
20 to create continuous, relatively high speed oscillations of the
air jets.
In an alternate embodiment of the present invention, the fuel
delivery is oscillated in form rather than in amount. By this it is
meant that the means for delivering fuel to the front end of the
combustion chamber automatically changes the fuel spray pattern to
cause an oscillating heat release rate. In the can type burner
shown in FIGS. 1 through 3 the size of the recirculation zone and
the amount of heat released by it can be modified by modifying the
spray pattern of the fuel. FIG. 4 shows the front end of a
combustion chamber similar to the combustion chambers of FIGS. 1
through 3 and having a fuel nozzle 52 (in place of the fuel nozzles
32) designed for oscillation of the fuel spray pattern. The nozzle
52 is surrounded by swirl vanes 54 and is disposed in the front end
of a burner can 56. The can 56 includes a row of primary combustion
air holes (not shown) similar to the holes 38 of FIG. 1. The nozzle
52 includes a centerline 62 coincident with the burner can
centerline.
Referring now to FIGS. 4 and 5, the nozzle 52 includes an outer
annular conical passage 64 and an inner annular conical passage 66
which are in communication at a common apex area 68. The nozzle 52
also includes inner and outer feedback loops 70, 72, respectively.
The downstream end 74 of the inner loop 70 is in communication with
the downstream end 76 of the inner annular passage 66 and the
upstream end 78 of the inner loop 70 is in communication with the
apex area 68 and is located a short distance upstream of the
upstream end 80 of the inner passage 66.
Similarly, the downstream end 82 of the outer loop 72 is in
communication with the downstream end 84 of the outer passage 64
while the upstream end 86 of the outer loop 72 is in communication
with the apex area 68 and is positioned a short distance upstream
of the upstream end 88 of the outer passage 64.
The nozzle 52 operates in the following manner: Fuel or a fuel-air
mixture from a suitable source represented by the box 89 is
supplied into the apex area 68. As the fuel flows into the inner
passage 66, it causes a sudden pressure buildup at the downstream
end 74 of the inner feedback loop 70. This high pressure in the
inner feedback loop 70 travels as a pulse to the upstream end 78 of
the inner feedback loop 70 whereupon it impinges upon the fuel
stream and diverts it to the outer passage 64. The fuel then
travels downstream in the outer passage 64 building up a high
pressure at the downstream end 82 of the outer feedback loop 72.
The high pressure in the outer feedback loop 72 travels as a pulse
to the upstream end 86 of the outer feedback loop 72, whereupon it
impinges upon the fuel stream and diverts it back to the inner
passage 66. The nozzle 52 is thus bistable, causing the fuel to
flip back and forth between the inner and outer passages 66, 64
respectively. In this manner the fuel spray oscillates between a
wide cone angle .theta..sub.2 and a narrow cone angle
.theta..sub.1.
As fuel is sprayed into the primary combustion zone of the outer
passage 64, its high radial momentum leads to a high concentration
of fuel in the primary combustion zone followed by an intense heat
release and high pressure buildup. When the fuel is sprayed through
the inner passage 66, its low radial momentum reduces the
concentration of fuel in the primary combustion zone; in that
instance the heat release rate is less intense and the pressure
buildup is not as great. The fuel oscillations thus amplify the
swings in pressure which occur in prior art combustion
chambers.
Merely oscillating the fuel spray pattern in a gas turbine engine
combustion chamber is old as shown in U.S. Pat. No. 3,039,699 to
Allen. The present invention teaches coupling the frequency of the
oscillations of the fuel spray pattern with the natural combustion
chamber acoustics. In that way the oscillations of the air jets
entering the primary combustion air holes are controlled and
amplified rather than being random. Further, as hereinabove
explained, it is desired that the frequency of oscillations of the
air jets be high enough so that the period of oscillation is
shorter than the vane thermal response time. The frequency of
oscillation of the fuel spray pattern is easily set by proper
choice of the length and the size of the feedback loops 70, 72. The
Allen patent does not teach this embodiment of the present
invention it merely teaches continuously varying the fuel spray
pattern so that "the individual particles will be more widely
dispersed and more evenly distributed." (Column 2, lines
44-46).
Another embodiment of a bistable nozzle is shown in FIGS. 6 and 7.
The nozzle, generally represented by the numeral 100, is quite
similar in operation and construction to the nozzle of FIGS. 4 and
5. The nozzle 100 includes inner and outer annular passages 102,
104 respectively, having their respective downstream ends 106, 108
configured to spray fuel in conical patterns. The inner passage 102
sprays fuel in the form of a cone having been included angle
.alpha..sub.1 while the outer passage 104 sprays fuel in the form
of a cone having an included angle .alpha..sub.2. The upstream ends
110, 112 of the inner and outer passages 102, 104, respectively,
are in communication with each other at a common annular apex area
114. Disposed within each passage 102, 104 are a plurality of
circumferentially spaced swirl vanes 116 which are configured to
provide, in combination with the configuration of the downstream
ends 106, 108, the desired conical spray angles .alpha..sub.1,
.alpha..sub.2. The nozzle 100 also includes a plurality of outer
feedback loops 118 and inner feedback loops 120. The outer feedback
loops 118 differ from the outer feedback loop 72 of the embodiment
shown in FIG. 4 in that the feedback loops 118 are discreet
cylindrical passageways as best shown in FIG. 7. Similarly,
although the inner feedback loop 120 includes a single cylindrical
axial passage 122, the pressure pulses are transmitted by means of
individual cylindrical radial passages 124 extending outwardly from
both the upstream and downstream end of the passage 122.
Fuel from a suitable source represented by the box 126 is fed into
the common apex area 114 and from there passes alternately through
the inner and outer annular passages 102, 104 whereupon it is
sprayed into the primary combustion zone. Although the construction
of the nozzle 100 is somewhat different in appearance than the
construction of the nozzle 52 of FIGS. 4 and 5, it operates in
almost identical fashion and further explanation of the operation
of the nozzle 100 is not deemed to be necessary.
FIG. 8 shows a schematic representation of another embodiment of
the present invention. A nozzle 150 positioned along the centerline
152 of a burner can 154, includes an inner annular conical fuel
spray passage 156 communicating with a fuel supply line 158 by
means of a central cylindrical passageway 160. The nozzle 150 also
includes an outer annular conical passage 162 fed from a fuel line
164 by means of an annular cavity 166. The fuel lines 158, 164 are
connected to a diverter valve 168 which is fed from a suitable fuel
source 170. The diverter valve 116 is connected to a timer 172. The
embodiment of FIG. 8 is intended to perform the identical functions
of the embodiment of FIGS. 4 and 5 except that the inner and outer
passages 156, 162, respectively, have separate fuel supplies. The
operation is quite simple. The timer 172 controls the diverter
valve 168 to alternate the fuel flow between the fuel lines 158,
164. The timer is of an adjustable type and can be set to provide
any frequency of oscillation of the fuel spray desired.
FIGS. 9 and 10 depict yet another embodiment of the present
invention, particularly adapted for use in an annular combustion
chamber. FIG. 9 shows an annular combustion chamber generally
represented by the numeral 200. The chamber 200 comprises inner and
outer cylindrical casing 202, 204, respectively, forming an annular
compartment 206 within which is an annular burner can generally
represented by the numeral 208. The can 208 comprises inner and
outer cylindrical liners 210, 212 respectively forming an annular
combustion zone 214. A plurality of circumferentially spaced fuel
nozzles 216 are positioned at the upstream end of the can 210 for
spraying fuel into the combustion zone 214. As best shown in FIG.
10, the inner and outer liners 210, 212 include a plurality of
circumferentially spaced combustion air holes 218, 220,
respectively, a short distance downstream of the fuel nozzles 216.
Although not shown, there are additional rows of holes similar to
the holes 218, 220 along the length of the can 208.
For reasons which have hereinbefore been discussed, it is desirable
to cause the air jets entering the holes 218, 220, and other air
jets entering the combustion zone 214 to oscillate at a steady,
high frequency so that the air jets sweep across the turbine inlet
guide vanes positioned at the downstream end of the combustion zone
214. This sweeping action tends to reduce hot spots caused by
concentrated prolonged impingment of an air jet at one location on
a vane. In an annular combustion chamber, it is more difficult to
establish longitudinal oscillations and longitudinal waves of
temperature within the combustion zone 214 since these waves and
oscillations tend to dissipate in a circumferential direction.
Advantage is taken of this fact by forcing the air jets, in this
embodiment, to oscillate circumferentially at a steady, high
frequency. This is accomplished by varying the fuel flow through
the nozzles 216 so that alternate nozzles have a high fuel flow at
the time that adjacent nozzles have a low fuel flow. This will
result in alternate zones of high and low pressure around the
upstream end of the combustion zone 214. The motion of the air jets
entering the holes 218, 220 is represented by the solid arrows 222
and the dotted arrows 224. The air jets will tend to move toward
the low pressure zones. Thus, as the low pressure zones alternate
back and forth between adjacent nozzles, so the air jets will swing
back and forth between the solid and dotted positions 222, 224.
One method for doing this is shown in FIG. 9 wherein two fuel
manifolds 226, 228 are used. The nozzles 216 are alternately
connected to the manifolds 226, 228. Fuel from a suitable source
represented by the box 230 is fed into a valve 232 which alternates
a high and low fuel flow between the manifolds 226, 228. For
example, the flow of fuel into each manifold 226, 228 might be
represented by a sine wave, wherein the fuel flow is measured along
the vertical axis and time is measured along the horizontal axis;
the sine wave for the fuel flowing through the manifold 226 should
be 180.degree. out of phase with the sine wave of the fuel flowing
through the manifold 228. In that way when one nozzle is flowing at
its maximum rate the nozzles on either side of it will be flowing
at their minimum rate.
Although the invention has been shown and described with respect to
preferred embodiments thereof, it should be understood by those
skilled in the art that various changes and omissions in the form
and detail thereof may be made therein without departing from the
spirit and the scope of the invention.
* * * * *