US20070028618A1

US20070028618A1 - Mixer assembly for combustor of a gas turbine engine having a main mixer with improved fuel penetration

Info

Publication number: US20070028618A1
Application number: US11/188,598
Authority: US
Inventors: George Hsiao; Shih-yang Hsieh; Shui-Chi Li; Hukam Mongia; Marie McMasters
Original assignee: General Electric Co
Current assignee: General Electric Co
Priority date: 2005-07-25
Filing date: 2005-07-25
Publication date: 2007-02-08

Abstract

A mixer assembly for use in a combustion chamber of a gas turbine engine includes a pilot mixer, a main mixer, and a fuel manifold positioned between the pilot mixer and main mixer. The pilot mixer includes an annular pilot housing having a hollow interior and a pilot fuel nozzle mounted in the housing and adapted for dispensing droplets of fuel to the hollow interior of the pilot housing. The main mixer includes: a main housing surrounding the pilot housing and defining an annular cavity having an upstream end and a downstream end including an upstream wall, an outer wall and an inner wall; a plurality of fuel injection ports for introducing fuel into the cavity, with the fuel injection ports being circumferentially spaced at a designated axial location of the inner wall of the annular cavity; and a swirler arrangement including at least one swirler in flow communication with the annular cavity, the swirler being incorporated into the outer wall of the annular cavity and extending from an upstream end to a downstream end, wherein each swirler of the arrangement has a plurality of vanes for swirling air traveling through such swirler to mix air and the droplets of fuel dispensed by the fuel injection ports. The main housing further includes a first plurality of passages oriented to provide air jets in a substantially axial direction into the annular cavity and a second plurality of passages oriented to provide air jets in a substantially radial direction into the annular cavity

Description

BACKGROUND OF THE INVENTION

The present invention relates to a staged combustion system in which the production of undesirable combustion product components is minimized over the engine operating regime and, more particularly, to a fuel injection arrangement for the main mixer of such system which enhances fuel penetration into an annular cavity for improved mixing of fuel and air therein.
Air pollution concerns worldwide have led to stricter emissions standards both domestically and internationally. Aircraft are governed by both Environmental Protection Agency (EPA) and International Civil Aviation Organization (ICAO) standards. These standards regulate the emission of oxides of nitrogen (NOx), unburned hydrocarbons (HC), and carbon monoxide (CO) from aircraft in the vicinity of airports, where they contribute to urban photochemical smog problems. Such standards are driving the design of gas turbine engine combustors, which also must be able to accommodate the desire for efficient, low cost operation and reduced fuel consumption. In addition, the engine output must be maintained or even increased.
It will be appreciated that engine emissions generally fall into two classes: those formed because of high flame temperatures (NOx) and those formed because of low flame temperatures which do not allow the fuel-air reaction to proceed to completion (HC and CO). Balancing the operation of a combustor to allow efficient thermal operation of the engine, while simultaneously minimizing the production of undesirable combustion products, is difficult to achieve. In that regard, operating at low combustion temperatures to lower the emissions of NOx can also result in incomplete or partially incomplete combustion, which can lead to the production of excessive amounts of HC and CO, as well as lower power output and lower thermal efficiency. High combustion temperature, on the other hand, improves thermal efficiency and lowers the amount of HC and CO, but oftentimes results in a higher output of NOx.
One way of minimizing the emission of undesirable gas turbine engine combustion products has been through staged combustion. In such an arrangement, the combustor is provided with a first stage burner for low speed and low power conditions so the character of the combustion products is more closely controlled. A combination of first and second stage burners is provided for higher power output conditions, which attempts to maintain the combustion products within the emissions limits.
Another way that has been proposed to minimize the production of such undesirable combustion product components is to provide for more effective intermixing of the injected fuel and the combustion air. In this way, burning occurs uniformly over the entire mixture and reduces the level of HC and CO that results from incomplete combustion. While numerous mixer designs have been proposed over the years to improve the mixing of the fuel and air, improvement in the levels of undesirable NOx formed under high power conditions (i.e., when the flame temperatures are high) is still desired.
One mixer design that has been utilized is known as a twin annular premixing swirler (TAPS), which is disclosed in the following U.S. Pat. Nos. 6,354,072; 6,363,726; 6,367,262; 6,381,964; 6,389,815; 6,418,726; 6,453,660; 6,484,489; and, 6,865,889. Published U.S. patent application 2002/0178732 also depicts certain embodiments of the TAPS mixer. It will be understood that the TAPS mixer assembly includes a pilot mixer which is supplied with fuel during the entire engine operating cycle and a main mixer which is supplied with fuel only during increased power conditions of the engine operating cycle. Because improvements in NOx emissions during high power conditions are of current primary concern, modification of the main mixer in the assembly is needed to maximize fuel-air mixing therein.
As shown in the '964 and '815 patents, fuel is injected from a fuel manifold into the main mixer by means of a plurality of fuel injection ports. Such ports are generally located downstream of a ramp portion defining an inner radial surface of an annular cavity. It has been found that fuel injected into such annular cavity has a tendency to break apart more quickly than desired and/or reside too closely to the inner radial surface thereof. In either event, the ability of the fuel and air in the annular cavity to form a more uniform mixture is impeded.
Accordingly, there is a desire for a gas turbine engine combustor in which the production of undesirable combustion product components is minimized over a wide range of engine operating conditions. More specifically, a mixer assembly for such gas turbine engine combustor is desired which provides increased mixing of fuel and air so as to create a more uniform mixture. It is desired that the fuel spray be injected further into the annular cavity of the main mixer and that a flow field be created therein which is conducive to retarding break-up of the fuel spray.

BRIEF SUMMARY OF THE INVENTION

In a first exemplary embodiment of the invention, a mixer assembly for use in a combustor of a gas turbine engine is disclosed as including a pilot mixer, a main mixer, and a fuel manifold positioned between the pilot mixer and main mixer. The pilot mixer includes an annular pilot housing having a hollow interior and a pilot fuel nozzle mounted in the housing and adapted for dispensing droplets of fuel to the hollow interior of the pilot housing. The main mixer includes: a main housing surrounding the pilot housing and defining an annular cavity having an upstream end and a downstream end, with the annular cavity including an upstream wall, an outer wall and an inner wall; a plurality of fuel injection ports for introducing fuel into the cavity, with the fuel injection ports being circumferentially spaced at a designated axial location of the inner wall of the annular cavity; and, a swirler arrangement including at least one swirler in flow communication with the annular cavity, the swirler being incorporated into the outer wall of the annular cavity and extending from an upstream end to a downstream end, wherein each swirler of the arrangement has a plurality of vanes for swirling air traveling through such swirler to mix air and the droplets of fuel dispensed by the fuel injection ports. The fuel injection ports for introducing fuel into the main mixer cavity are in flow communication with the fuel manifold.
In a second exemplary embodiment of the invention, a mixer assembly for use in a combustor of a gas turbine engine is disclosed at including a pilot mixer, a main mixer and a fuel manifold positioned between the pilot mixer and the main mixer. The pilot mixer includes an annular pilot housing having a hollow interior and a pilot fuel nozzle mounted in the housing and adapted for dispensing droplets of fuel to the hollow interior of the pilot housing. The main mixer includes: a main housing surrounding the pilot housing and defining an annular cavity; a plurality of fuel injection ports for introducing fuel into the annular cavity; and, a swirler arrangement including at least one swirler positioned upstream from the fuel injection ports, wherein each swirler of the arrangement has a plurality of vanes for swirling air traveling through such swirler to mix air and the droplets of fuel dispensed by the fuel injection ports. The main housing of the main mixer further includes: a ramp portion positioned at an upstream portion of the annular cavity; an upstream wall including a first plurality of openings in flow communication with an air supply, where the first openings are oriented to provide air jets in a substantially axial direction into the annular cavity; and, an axial wall downstream of the upstream wall including a second plurality of openings in flow communication with an air supply oriented to provide air jets in a substantially radial direction into the annular cavity. The fuel injection ports are positioned adjacent the ramp portion of the annular cavity and are in flow communication with the fuel manifold.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a diagrammatic view of a high bypass turbofan gas turbine engine;
FIG. 2 is a longitudinal, cross-sectional view of a gas turbine engine combustor having a staged arrangement;
FIG. 3 is an enlarged, cross-sectional view of the mixer assembly depicted in FIG. 2;
FIG. 4 is a partial perspective view of the mixer assembly depicted in FIGS. 2 and 3;
FIG. 5 is a front perspective view of the swirler arrangement removed from the mixer assembly depicted in FIGS. 2-4;
FIG. 6 is an aft perspective view of the swirler arrangement depicted in FIG. 5, where a portion thereof has been removed for clarity;
FIG. 7 is an exemplary coordinate system provided as a reference for the orientation of openings depicted in the swirler arrangement of FIGS. 5 and 6;
FIG. 8 is a partial rear view of a portion of the swirler arrangement depicted in FIGS. 5 and 6, where openings formed in an upstream wall are shown in greater detail;
FIG. 9 is a partial side view of a portion of the swirler arrangement depicted in FIGS. 5 and 6, where openings formed in an inner wall are shown in detail;
FIG. 10 is a partial section view of the swirler arrangement depicted in FIGS. 5 and 6, where the orientation of the openings depicted in FIG. 8 are shown therein;
FIG. 11 is a partial perspective view of the mixer assembly depicted in FIGS. 3 and 4, where the swirler arrangement has been removed for clarity;
FIG. 12 is a partial perspective view of the mixer assembly similar to that depicted in FIG. 11, where the fuel injection ports are located downstream of the ramp portion in the main mixer and openings are formed in the axial wall upstream of the ramp portion;
FIG. 13 is a partial perspective view of the mixer assembly similar to that depicted in FIG. 12, where a row of purge holes is provided upstream of the fuel injection ports;
FIG. 14 is a partial perspective view of the mixer assembly similar to that depicted in FIG. 12, where a row of purge holes is provided downstream of the fuel injection ports;
FIG. 15 is an enlarged, cross-sectional view of the mixer assembly including a second embodiment for the swirler arrangement depicted in FIGS. 3-6 and 8-10, where a radial swirler and an axial swirler are provided;
FIG. 16 is an enlarged, partial perspective view of the mixer assembly depicted in FIG. 15;
FIG. 17 is an enlarged, cross-sectional view of the mixer assembly including a third embodiment for the swirler arrangement depicted in FIGS. 3-6 and 8-10, where a conical swirler and an axial swirler are provided; and,
FIG. 18 is an enlarged, partial perspective view of the mixer assembly depicted in FIG. 17.

DETAILED DESCRIPTION OF THE INVENTION

Referring now to the drawings in detail, wherein identical numerals indicate the same elements throughout the figures, FIG. 1 depicts in diagrammatic form an exemplary gas turbine engine 10 (high bypass type) utilized with aircraft having a longitudinal or axial centerline axis 12 therethrough for reference purposes. Engine 10 preferably includes a core gas turbine engine generally identified by numeral 14 and a fan section 16 positioned upstream thereof. Core engine 14 typically includes a generally tubular outer casing 18 that defines an annular inlet 20. Outer casing 18 further encloses and supports a booster compressor 22 for raising the pressure of the air that enters core engine 14 to a first pressure level. A high pressure, multi-stage, axial-flow compressor 24 receives pressurized air from booster 22 and further increases the pressure of the air. The pressurized air flows to a combustor 26, where fuel is injected into the pressurized air stream to raise the temperature and energy level of the pressurized air. The high energy combustion products flow from combustor 26 to a first (high pressure) turbine 28 for driving high pressure compressor 24 through a first (high pressure) drive shaft 30, and then to a second (low pressure) turbine 32 for driving booster compressor 22 and fan section 16 through a second (low pressure) drive shaft 34 that is coaxial with first drive shaft 30. After driving each of turbines 28 and 32, the combustion products leave core engine 14 through an exhaust nozzle 36 to provide propulsive jet thrust.
Fan section 16 includes a rotatable, axial-flow fan rotor 38 that is surrounded by an annular fan casing 40. It will be appreciated that fan casing 40 is supported from core engine 14 by a plurality of substantially radially-extending, circumferentially-spaced outlet guide vanes 42. In this way, fan casing 40 encloses fan rotor 38 and fan rotor blades 44. Downstream section 46 of fan casing 40 extends over an outer portion of core engine 14 to define a secondary, or bypass, airflow conduit 48 that provides additional propulsive jet thrust.
From a flow standpoint, it will be appreciated that an initial air flow, represented by arrow 50, enters gas turbine engine 10 through an inlet 52 to fan casing 40. Air flow 50 passes through fan blades 44 and splits into a first compressed air flow (represented by arrow 54) that moves through conduit 48 and a second compressed air flow (represented by arrow 56) which enters booster compressor 22. The pressure of second compressed air flow 56 is increased and enters high pressure compressor 24, as represented by arrow 58. After mixing with fuel and being combusted in combustor 26, combustion products 60 exit combustor 26 and flow through first turbine 28. Combustion products 60 then flow through second turbine 32 and exit exhaust nozzle 36 to provide thrust for gas turbine engine 10.
As best seen in FIG. 2, combustor 26 includes an annular combustion chamber 62 that is coaxial with longitudinal axis 12, as well as an inlet 64 and an outlet 66. As noted above, combustor 26 receives an annular stream of pressurized air from a high pressure compressor discharge outlet 69. A portion of this compressor discharge air flows into a mixing assembly 67, where fuel is also injected from a fuel nozzle 68 to mix with the air and form a fuel-air mixture that is provided to combustion chamber 62 for combustion. Ignition of the fuel-air mixture is accomplished by a suitable igniter 70, and the resulting combustion gases 60 flow in an axial direction toward and into an annular, first stage turbine nozzle 72. Nozzle 72 is defined by an annular flow channel that includes a plurality of radially-extending, circularly-spaced nozzle vanes 74 that tun the gases so that they flow angularly and impinge upon the first stage turbine blades of first turbine 28. As shown in FIG. 1, first turbine 28 preferably rotates high pressure compressor 24 via first drive shaft 30. Low pressure turbine 32 preferably drives booster compressor 24 and fan rotor 38 via second drive shaft 34.
Combustion chamber 62 is housed within engine outer casing 18 and is defined by an annular combustor outer liner 76 and a radially-inwardly positioned annular combustor inner liner 78. The arrows in FIG. 2 show the directions in which compressor discharge air flows within combustor 26. As shown, part of the air flows over the outermost surface of outer liner 76, part flows into combustion chamber 62, and part flows over the innermost surface of inner liner 78.
Contrary to previous designs, it is preferred that outer and inner liners 76 and 78, respectively, not be provided with a plurality of dilution openings to allow additional air to enter combustion chamber 62 for completion of the combustion process before the combustion products enter turbine nozzle 72. This is in accordance with a patent application entitled “Combustor Liner Having No Dilution Holes,” filed concurrently herewith and hereby incorporated by reference, which is also owned by the assignee of the present invention. It will be understood, however, that outer liner 76 and inner liner 78 preferably include a plurality of smaller, circularly-spaced cooling air apertures (not shown) for allowing some of the air that flows along the outermost surfaces thereof to flow into the interior of combustion chamber 62. Those inwardly-directed air flows pass along the inner surfaces of outer and inner liners 76 and 78 that face the interior of combustion chamber 62 so that a film of cooling air is provided therealong.
It will be understood that a plurality of axially-extending mixing assemblies 67 are disposed in a circular array at the upstream end of combustor 26 and extend into inlet 64 of annular combustion chamber 62. It will be seen that an annular dome plate 80 extends inwardly and forwardly to define an upstream end of combustion chamber 62 and has a plurality of circumferentially spaced openings formed therein for receiving mixing assemblies 67. For their part, upstream portions of each of inner and outer liners 76 and 78, respectively, are spaced from each other in a radial direction and define an outer cowl 82 and an inner cowl 84. The spacing between the forwardmost ends of outer and inner cowls 82 and 84 defines combustion chamber inlet 64 to provide an opening to allow compressor discharge air to enter combustion chamber 62.
A mixing assembly 100 in accordance with one embodiment of the present invention is shown in FIG. 3. Mixing assembly 100 preferably includes a pilot mixer 102, a main mixer 104, and a fuel manifold 106 positioned therebetween. More specifically, it will be seen that pilot mixer 102 preferably includes an annular pilot housing 108 having a hollow interior, as well as a pilot fuel nozzle 110 mounted in housing 108 and adapted for dispensing droplets of fuel to the hollow interior of pilot housing 108. Further, pilot mixer 102 preferably includes a first swirler 112 located at a radially inner position adjacent pilot fuel nozzle 110, a second swirler 114 located at a radially outer position from first swirler 112, and a splitter 116 positioned therebetween. Splitter 116 extends downstream of pilot fuel nozzle 110 to form a venturi 118 at a downstream portion. It will be understood that first and second pilot swirlers 112 and 114 are generally oriented parallel to a centerline axis 120 through mixing assembly 100 and include a plurality of vanes for swirling air traveling therethrough. Fuel and air are provided to pilot mixer 102 at all times during the engine operating cycle so that a primary combustion zone 122 is produced within a central portion of combustion chamber 62 (see FIG. 2).
Main mixer 104 further includes an annular main housing 124 radially surrounding pilot housing 108 and defining an annular cavity 126, a plurality of fuel injection ports 128 which introduce fuel into annular cavity 126, and a swirler arrangement identified generally by numeral 130. More specifically, annular cavity 126 is preferably defined by an upstream wall 132 and an outer radial wall 134 of a swirler housing 136, and by an inner radial wall 138 of a centerbody outer shell 140. It will be seen that inner radial wall 138 preferably also includes a ramp portion 142 located at a forward position along annular cavity 126. It will be appreciated that annular cavity 126 gently transitions from an upstream end 127 having a first radial height 129 to a downstream end 131 having a second radial height 133. The difference between first radial height 129 and second radial height 133 of annular cavity 126 is due primarily to outer radial wall 134 of swirler housing 136 incorporating a swirler 144 therein at upstream end 127. In addition, ramp portion 142 of inner radial wall 138 is preferably located within an axial length 145 of swirler 144.
It will be seen in FIGS. 3-6 and 10 that swirler arrangement 130 preferably includes at least a first swirler 144 positioned upstream from fuel injection ports 128. As shown, first swirler 144 is preferably oriented substantially radially to centerline axis 120 through mixer assembly 100 and has an axis 148 therethrough. It will be noted that first swirler 144 includes a plurality of vanes 150 extending between first and second portions 137 and 139 of outer radial wall 134. It will be appreciated that vanes 150 are preferably oriented at an angle of approximately 30-70° with respect to axis 148. Vanes 150 will preferably each have a height 151 which is measured across opposite ends (i.e., in the axial direction relative to centerline axis 120 of mixing assembly 100) that is equivalent to axial length 145 of swirler 144. Since vanes 150 are substantially uniformly spaced circumferentially, a plurality of substantially uniform passages 154 are defined between adjacent vanes 150. It will be noted that vanes 150 preferably extend from upstream end 147 of swirler 144 to downstream end 149 thereof. Nevertheless, vanes 150 may extend only part of the way from upstream end 147 to downstream end 149 so that the tips thereof are stepped or lie on a different annulus. It will further be understood that swirler 144 may include vanes having different configurations so as to shape the passages in a desirable manner, as disclosed in a patent application entitled “Swirler Arrangement For Mixer Assembly Of A Gas Turbine Engine Combustor Having Shaped Passages,” which is also filed concurrently herewith by the assignee of the present invention and is hereby incorporated herein.
Air is also provided at upstream end 127 of annular cavity 126 via a series of passages formed in upstream wall 132 of swirler housing 130. More specifically, as best seen in FIGS. 4-6, an outer row 152 of passages 153 and an inner row 156 of passages 157 are provided which direct jets of air in a substantially axial direction into annular cavity 126. A third row 158 of passages 159 located within upstream wall 132 is further provided which direct jets of air in a substantially radial direction into annular cavity 126. It will be appreciated from FIGS. 7-10 that passages 153 preferably are oriented to be at an angle 155 in a range of approximately 30-70° in relation to centerline axis 120 relative to a radial plane 161. Passages 153 are also preferably oriented to be at an angle 160 in a range of approximately 30-70° with respect to a radial axis 162 relative to a tangential plane 163. Similarly, passages 157 preferably are oriented to be at an angle 164 in a range of approximately 30-70° with respect to centerline axis 120 relative to radial plane 161 and at an angle 165 in a range of approximately 30-70° with respect to radial axis 162 relative to tangential plane 163. It will be seen that passages 159 of third row 158 are preferably oriented to be at an angle 166 with respect to centerline axis 120 relative to radial plane 161, as well as at an angle 167 with respect to a radial axis 162 relative to tangential plane 163. Further, passages 153, 157 and 159 may be oriented at angles 173, 175 and 177, respectively, to a tangential axis 169 relative to an axial plane 171.
It will be understood that air flowing through swirler 144 will be swirled in a first direction and air flowing through passages 153, 157 and 159 will preferably be swirled in a direction opposite the first direction. In this way, an intense mixing region 168 of air and fuel is created within annular cavity 126 having an enhanced total kinetic energy. By properly configuring swirler 144, as well as passages 153, 157 and 159, intense mixing region 168 is substantially centered within annular cavity 126, positioned axially adjacent fuel injection ports 128 and has a designated area. The configuration of vanes 150 in swirler 144 and orientation of passages 153, 157 and 159 may be altered to vary the swirl direction of air flowing therethrough and not be limited to the exemplary swirl directions indicated hereinabove.
It will be understood that passages 154 between swirler vanes 150 preferably have a greater area than the cumulative area of passages 153, 157 and 159. Accordingly, a relatively greater amount of air flows through first swirler 144 than through passages 153, 157 and 159 due to the greater passage area therefor. The relative area of swirler passages 154 and passages 153, 157 and 159, however, may be varied as desired to alter the distribution of air therethrough, so the sizes depicted are only illustrative. Regarding the amount of air flowing through passages 153, 157 and 159, it is preferred that this be approximately 15-30% of the total air flowing through main mixer 104.
Fuel manifold 106, as stated above, is located between pilot mixer 102 and main mixer 104 and is in flow communication with a fuel supply. In particular, outer radial wall 138 of centerbody outer shell 140 forms an outer surface 170 of fuel manifold 106, and a shroud member 172 is configured to provide an inner surface 174 and an aft surface 176 thereof. Fuel injection ports 128 are in flow communication with fuel manifold 106 and spaced circumferentially around centerbody outer shell 140. As seen in FIG. 3, fuel injection ports 128 are preferably positioned axially adjacent ramp portion 142 of centerbody outer shell 140 so that fuel is provided in upstream end 127 of annular cavity 126. In this way, fuel is preferably mixed with the air in intense mixing region 168 before entering downstream end 131 of annular cavity 126. While fuel injection ports 128 are positioned upstream of ramp portion 142 in FIG. 3, it will be noted that such ports 128 may be located immediately downstream thereof (see FIGS. 12-14). In either case, it is preferred that the axial location of fuel injection ports 128 facilitate injection of fuel at least a specified distance into a middle radial portion of annular cavity 126 and away from the surface of inner radial wall 138. Accordingly, such fuel injection ports 128 will generally be located within axial length 145 of radial swirler 144.
It will be appreciated that injection of the fuel into the desired location of annular cavity 126 is a function of providing an air flow therein which accommodates such injected fuel (instead of forcing the fuel against inner radial wall 138), as well as positioning fuel injection ports 128 so as to inject fuel in the manner best suited to the air flow. In addition, at least one row of circumferentially spaced purge holes is provided adjacent to and between each fuel injection port 128 to assist the injected fuel in its intended path. Such purge holes also assist in preventing injected fuel from collecting along inner radial wall 138. More specifically, it will be seen in FIGS. 3 and 4 that a first row of purge holes 179 is located immediately upstream of and between fuel injection ports 128, a second row of purge holes 180 is located immediately downstream of and between fuel injection ports 128, and third and fourth rows of purge holes 181 and 182 are located between adjacent fuel injection ports 128. An alternative configuration of purge holes may be utilized, as shown in FIG. 11, where a downstream row of purge holes 183 is located within ramp portion 142 further downstream from fuel injection ports 128 and only a single row of intermediate purge holes 184 is located between adjacent fuel injection ports 128.
Depending on the axial location of fuel injection ports 128, the location of its purge holes may also be altered. For example, rows of purge holes 185, 186, 187 and 188 are located upstream of ramp portion 142 when fuel injection ports 128 are located downstream of such ramp portion 142 (see FIGS. 12-14). Moreover, it will be noted that an additional row of purge holes 189 may be included upstream (FIG. 13) or downstream (FIG. 14) of fuel injection ports 128.
In order to further facilitate injection of the fuel from fuel injection ports 128 into annular cavity 126, it is also preferred that a post member 190 having an inner passage 191 be associated with each such fuel injection port 128. It will be seen that post member 190 preferably extends from fuel injection port 128 through an air cavity 192 supplying compressed air to all applicable purge holes discussed hereinabove and through inner wall 138. In this way, fuel not only is injected directly into annular cavity 126, but the fuel is better able to travel into a middle annular portion of annular cavity 126 with the assistance of purge holes 179, 180, 181 and 182.
When fuel is provided to main mixer 104, an annular, secondary combustion zone 178 is provided in combustion chamber 62 that is radially outwardly spaced from and concentrically surrounds primary combustion zone 122. Depending upon the size of gas turbine engine 10, as many as twenty or so mixer assemblies 100 can be disposed in a circular array at inlet 64 of combustion chamber 62.
FIGS. 15-16 depict an alternative swirler arrangement 200, where swirler housing 202 includes a first swirler 204 oriented substantially radially to centerline axis 120 similar to swirler 144 described hereinabove with an axis 205 therethrough. In addition, swirler housing 202 includes a second swirler 206 which is oriented substantially axially to centerline axis 120 and is utilized to provide the counter swirling flow in annular cavity 126 instead of passages 153, 157 and 159. It will be appreciated that first and second swirlers 204 and 206 will each preferably have a plurality of vanes 208 and 210, respectively, with passages 212 and 214 defined therebetween to provide the intended swirling air flows into annular cavity 126 to mix with the injected fuel.
It will be appreciated that vanes 208 of first swirler 204 are oriented at an angle of approximately 30-70° with respect to axis 205. Vanes 208 each have a length 216 which is measured across opposite ends (i.e., in the axial direction perpendicular to axis 205). Because vanes 208 are uniformly spaced circumferentially around swirler housing 202, passages 212 defined between adjacent vanes are uniform. It will be noted that vanes 208 preferably extend from an upstream end 218 of first swirler 204 to a downstream end 220. It will be understood, however, that first swirler 204 could include different vanes therein so as to form shaped passages therethrough.
Similarly, it will be appreciated that vanes 210 of second swirler 206 are oriented at an angle of approximately 30-70° with respect to an axis 222 parallel to centerline axis 120. Vanes 210 each have a length 224 which is measured across opposite ends (i.e., in the radial direction perpendicular to axis 222). Because vanes 210 are uniformly spaced circumferentially around swirler housing 202, passages 214 defined between adjacent vanes are uniform. It will be noted that vanes 210 preferably extend from an inner radial end 226 of second swirler 206 to an outer radial end 228. It will be understood that second swirler 206 could include different vanes therein so as to form shaped passages therethrough.
It will be understood that air flowing through first swirler 204 will be swirled in a first direction and air flowing through second swirler 206 will preferably be swirled in a direction opposite the first direction. In this way, an intense mixing region 230 of air and fuel is created within annular cavity 126 having an enhanced total kinetic energy. By properly configuring swirlers 204 and 206, intense mixing region 230 is substantially centered within annular cavity 126, positioned axially adjacent fuel injection ports 128 and has a designated area. The configuration of the vanes in swirlers 204 and 206 may be altered to vary the swirl direction of air flowing therethrough and not be limited to the exemplary swirl directions indicated hereinabove.
It will be seen that length 216 of first swirler vanes 208 is preferably greater than height 224 of second swirler vanes 210. Accordingly, a relatively greater amount of air flows through first swirler 204 than second swirler 206 due to the greater passage area therefor. The relative lengths of swirlers 204 and 206 may be varied as desired to alter the distribution of air flowing therethrough, so the sizes depicted are only illustrative.
FIGS. 17 and 18 depict another alternative swirler arrangement 300 having a swirler housing 302 where swirler housing 302 includes a first swirler 304 oriented at an acute angle 303 to centerline axis 120 with an axis 305 therethrough. In addition, swirler housing 302 includes a second swirler 306 which is oriented substantially axially to centerline axis 120 and is utilized to provide the counter swirling flow in annular cavity 126. It will be appreciated that first and second swirlers 304 and 306 will each preferably have a plurality of vanes 308 and 310, respectively, with passages 312 and 314 defined therebetween to provide the intended swirling air flows into annular cavity 126 to mix with the injected fuel.
It will be appreciated that vanes 308 of first swirler 304 are oriented at an angle of approximately 30-70° with respect to axis 305. Vanes 308 each have a length 316 which is measured across opposite ends (i.e., in the axial direction perpendicular to axis 305). Because vanes 308 are uniformly spaced circumferentially around swirler housing 302, passages 312 defined between adjacent vanes are uniform. It will be noted that vanes 308 preferably extend from an upstream end 318 of first swirler 304 to a downstream end 320. It will be understood that first swirler 304 could include different vanes therein so as to form shaped passages therethrough.
Similarly, it will be appreciated that vanes 310 of second swirler 306 are oriented at an angle of approximately 30-70° with respect to an axis 322 parallel to centerline axis 120. Vanes 310 each have a length 324 which is measured across opposite ends (i.e., in the radial direction perpendicular to axis 322). Because vanes 310 are uniformly spaced circumferentially around swirler housing 302, passages 314 defined between adjacent vanes are uniform. It will be noted that vanes 310 preferably extend from an inner radial end 326 of second swirler 306 to an outer radial end 328. It will be understood that second swirler 306 could include different vanes therein so as to form shaped passages therethrough.
It will be understood that air flowing through first swirler 304 will be swirled in a first direction and air flowing through second swirler 306 will preferably be swirled in a direction opposite the first direction. In this way, an intense mixing region 330 of air and fuel is created within annular cavity 126 having an enhanced total kinetic energy. By properly configuring swirlers 304 and 306, intense mixing region 330 is substantially centered within annular cavity 126, positioned axially adjacent fuel injection ports 128 and has a designated area. The configuration of the vanes in swirlers 304 and 306 may be altered to vary the swirl direction of air flowing therethrough and not be limited to the exemplary swirl directions indicated hereinabove.
It will be seen that length 316 of first swirler vanes 308 is preferably greater than length 324 of second swirler vanes 310. Accordingly, a relatively greater amount of air flows through first swirler 304 than second swirler 306 due to the greater passage area therefor. The relative lengths of swirlers 304 and 306 may be varied as desired to alter the distribution of air flowing therethrough, so the sizes depicted are only illustrative.
Although particular embodiments of the present invention have been illustrated and described, it will be apparent to those skilled in the art that various changes and modifications can be made without departing from the spirit of the present invention. Accordingly, it is intended to encompass within the appended claims all such changes and modification that fall within the scope of the present invention.

Claims

1. A mixer assembly for use in a combustion chamber of a gas turbine engine, comprising:

(a) a pilot mixer including an annular pilot housing having a hollow interior and a pilot fuel nozzle mounted in said housing and adapted for dispensing droplets of fuel to said hollow interior of said pilot housing;

(b) a main mixer including:

(1) a main housing surrounding said pilot housing and defining an annular cavity having an upstream end and a downstream end, said annular cavity including an upstream wall; an outer wall and an inner wall;

(2) a plurality of fuel injection ports for introducing fuel into said annular cavity, said fuel injection ports being circumferentially spaced at a designated axial location of said inner wall of said annular cavity; and,

(3) a swirler arrangement including at least one swirler in flow communication with said annular cavity, said swirler being incorporated into said outer wall of said annular cavity and extending from an upstream end to a downstream end, wherein each swirler of said arrangement has a plurality of vanes for swirling air traveling through such swirler to mix air and said droplets of fuel dispensed by said fuel injection ports; and,

(c) a fuel manifold positioned between said pilot mixer and said main mixer, wherein said plurality of fuel injection ports for introducing fuel into said main mixer cavity are in flow communication with said fuel manifold.

2. The mixer assembly of claim 1, said upstream wall of said annular cavity including a first plurality of circumferentially spaced passages in flow communication with compressed air.

3. The mixer assembly of claim 2, wherein said first passages are oriented to provide air jets in a substantially axial direction into said annular cavity.

4. The mixer assembly of claim 3, wherein said first passages are oriented to provide a designated swirl to said air jets in said annular cavity.

5. The mixer assembly of claim 4, wherein said designated swirl of said air jets is at an angle in a range of about 25° to about 60° with respect to a centerline axis through said mixer assembly.

6. The mixer assembly of claim 4, wherein said air jets are swirled in a direction opposite air swirled by said swirler.

7. The mixer assembly of claim 2, said upstream wall of said annular cavity further including a second plurality of passages in flow communication with compressed air.

8. The mixer assembly of claim 7, wherein said second passages are oriented to provide air jets in a substantially radial direction into said annular cavity.

9. The mixer assembly of claim 8, wherein said second passages are oriented to provide a designated swirl to said air jets in said annular cavity.

10. The mixer assembly of claim 9, wherein said designated swirl of said air jets is at an angle in a range of about 25° to about 60° with respect to an axis substantially perpendicular to a centerline axis through said mixer assembly.

11. The mixer assembly of claim 9, wherein said air jets are swirled in a direction opposite air swirled by said swirler.

12. The mixer assembly of claim 1, wherein said fuel injection ports are located immediately downstream of a ramp portion in said inner wall of said annular cavity.

13. The mixer assembly of claim 1, wherein said fuel injection ports are located immediately upstream of a ramp portion in said inner wall of said annular cavity.

14. The mixer assembly of claim 1, wherein said fuel injection ports are spaced circumferentially around said annular cavity.

15. The mixer assembly of claim 14, further comprising a plurality of purge holes between each said fuel injection port in flow communication with compressed air.

16. The mixer assembly of claim 1, said fuel injection ports further comprising a post member having an inner passage in flow communication therewith.

17. The mixer assembly of claim 16, wherein said post member of said fuel injection ports extends at least even with an outer surface of said inner wall of said annular cavity.

18. The mixer assembly of claim 1, said swirler arrangement further comprising at least one swirler oriented substantially radially to a centerline axis through said mixer assembly having a designated axial length.

19. The mixer assembly of claim 18, wherein said injection ports are located within said axial length of said swirler.

20. The mixer assembly of claim 1, said swirler arrangement further comprising at least one swirler oriented at an acute angle to a centerline axis through said mixer assembly.

21. The mixer assembly of claim 1, said swirler arrangement further comprising at least one swirler oriented substantially parallel to a centerline axis through said mixer assembly.

22. The mixer assembly of claim 1, wherein fuel droplets from said fuel injection ports are able to penetrate to a designated position within said annular cavity.

23. The mixer assembly of claim 7, wherein air injected through said first and second plurality of passages is in a range of about 15% to about 30% of total air flowing through said main mixer.

24. A mixer assembly for use in a combustion chamber of a gas turbine engine, comprising:

(b) a main mixer including:

(1) a main housing surrounding said pilot housing and defining an annular cavity, said main housing further comprising:

(a) a ramp portion positioned at an upstream portion of said annular cavity;

(b) an upstream wall including a first plurality of passages and a second plurality of passages in flow communication with an air supply, wherein said first passages are oriented to provide air jets in a substantially axial direction into said annular cavity and said second passages are oriented to provide air jets in a substantially radial direction into said annular cavity;

(2) a plurality of fuel injection ports for introducing fuel into said annular cavity, wherein said fuel injection ports are positioned adjacent said ramp portion of said annular cavity; and,

(3) a swirler arrangement including at least one swirler positioned upstream from said fuel injection ports, wherein each swirler of said arrangement has a plurality of vanes for swirling air traveling through such swirler to mix air and said droplets of fuel dispensed by said fuel injection ports; and,