US3851462A - Method for reducing turbine inlet guide vane temperatures - Google Patents

Method for reducing turbine inlet guide vane temperatures Download PDF

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US3851462A
US3851462A US37524873A US3851462A US 3851462 A US3851462 A US 3851462A US 37524873 A US37524873 A US 37524873A US 3851462 A US3851462 A US 3851462A
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fuel
burner
combustion chamber
oscillating
method
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A Vranos
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United Technologies Corp
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23RGENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/28Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23RGENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/04Air inlet arrangements
    • F23R3/06Arrangement of apertures along the flame tube
    • YGENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02TCLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00Aeronautics or air transport
    • Y02T50/60Efficient propulsion technologies
    • Y02T50/67Relevant aircraft propulsion technologies
    • Y02T50/675Enabling an increased combustion temperature by cooling

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.

Description

ited States Vranos [1 11 3,851,62 [4 1 Dec. 3, 1974 METHOD FOR REDUCING TURBINE INLET GUIDE VANE TEMPERATURES [75] Inventor:

[73] Assignee: United Aircraft Corporation, East Hartford, Conn.

[22] Filed: June 29, 1973 [21] Appl. No.: 375,248

Alexander Vranos, Rockville, Conn.

[52] U.S. Cl 60/3906, 60/3928 R, 60/3937, 60/3965, 60/3974 R [51] Int. Cl F02c 7/12, F020 7/22 [58] Field of Search..... 60/3928 R, 39.36, 39.09 R, 60/3982 R, 39.46, 39.37, 39.74 R, 39.74 B, 39.06, 39.65; 431/12 000 o o 0 00f 3,039,699 6/1962 Allen 60/3974 R 3,053,047 9/1962 Bodemuller 60/3928 R 3,418,805 12/1968 Barish et al. 60/3928 R 3,531,934 10/1970 Hope-Gill 60/3937 3,688,495 9/1972 Fehler et al 60/3974 R 3,748,852 7/1973 Cole et al 60/3974 R 3,763,650 lO/1973 Hussey et a1. 60/3974 B Primary ExaminerCarlton R. Croyle Assistant Examiner-Warren Olsen Attorney, Agent, or FirmStephen E. Revis [57 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.

4 Claims, 10 Drawing Figures PATENTEL 31974 2.851.462

SHEET s (If a METHOD FOR REDUCING TURBINE INLET GUIDE VANE TEMPERATURES BACKGROUND OF THE INVENTION I. 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 US. Pat. No. 3,1 11,931; 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 55 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 10l0 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 chamberlO 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 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 at tempts 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 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 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 intothe 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 thenormal 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 noz proper cycle of high and low heat release rates can be zles 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 havihg 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 andis 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 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 0 and a narrow cone angle 6 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 ofa cone having been included angle a, while the outer passage 104 sprays fuel in the form of a cone having an included angle 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 (1 a 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 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 horizontalaxis; the sine wave for the fuel flowing through the manifold 226 should be 180 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 1 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 re- 'ducing 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 5 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 thefuel delivery includes oscillating the spray pattern of the fuel with said period of oscillation and atsaid se- 5 Iected 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.

Claims (4)

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.
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US4653278A (en) * 1985-08-23 1987-03-31 General Electric Company Gas turbine engine carburetor
US5490389A (en) * 1991-06-07 1996-02-13 Rolls-Royce Plc Combustor having enhanced weak extinction characteristics for a gas turbine engine
US6530228B2 (en) * 2001-05-07 2003-03-11 The United States Of America As Represented By The Secretary Of The Navy Method and device for modulation of a flame
US6640549B1 (en) * 2002-12-03 2003-11-04 The United States Of America As Represented By The Secretary Of The Navy Method and device for modulation of a flame
US20100043441A1 (en) * 2008-08-25 2010-02-25 William Kirk Hessler Method and apparatus for assembling gas turbine engines
US20100293953A1 (en) * 2007-11-02 2010-11-25 Siemens Aktiengesellschaft Combustor for a gas-turbine engine
US20130276450A1 (en) * 2012-04-24 2013-10-24 General Electric Company Combustor apparatus for stoichiometric combustion
US20160363041A1 (en) * 2015-06-15 2016-12-15 Caterpillar Inc. Combustion Pre-Chamber Assembly Including Fluidic Oscillator

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WO1979000773A1 (en) * 1978-03-16 1979-10-04 Caterpillar Tractor Co Dual phase fuel vaporizing combustor
US4242863A (en) * 1978-03-16 1981-01-06 Caterpillar Tractor Co. Dual phase fuel vaporizing combustor
US4653278A (en) * 1985-08-23 1987-03-31 General Electric Company Gas turbine engine carburetor
US5490389A (en) * 1991-06-07 1996-02-13 Rolls-Royce Plc Combustor having enhanced weak extinction characteristics for a gas turbine engine
US6530228B2 (en) * 2001-05-07 2003-03-11 The United States Of America As Represented By The Secretary Of The Navy Method and device for modulation of a flame
US6601393B2 (en) * 2001-05-07 2003-08-05 The United States Of America As Represented By The Secretary Of The Navy Method and device for modulation of a flame
US6640549B1 (en) * 2002-12-03 2003-11-04 The United States Of America As Represented By The Secretary Of The Navy Method and device for modulation of a flame
US20100293953A1 (en) * 2007-11-02 2010-11-25 Siemens Aktiengesellschaft Combustor for a gas-turbine engine
US8984889B2 (en) * 2007-11-02 2015-03-24 Siemens Aktiengesellschaft Combustor for a gas-turbine engine with angled pilot fuel nozzle
US20100043441A1 (en) * 2008-08-25 2010-02-25 William Kirk Hessler Method and apparatus for assembling gas turbine engines
US8397512B2 (en) 2008-08-25 2013-03-19 General Electric Company Flow device for turbine engine and method of assembling same
US20130276450A1 (en) * 2012-04-24 2013-10-24 General Electric Company Combustor apparatus for stoichiometric combustion
US20160363041A1 (en) * 2015-06-15 2016-12-15 Caterpillar Inc. Combustion Pre-Chamber Assembly Including Fluidic Oscillator

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