Patent Application
20070028618
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.
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.
Referring now to the drawings in detail, wherein identical numerals indicate the same elements throughout the figures,
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
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
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
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
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
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
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
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
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.
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.
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.
