Patent Application
20100251719
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 centerbody between the main mixer and the pilot mixer which includes an air manifold for supplying air to the main mixer and the pilot mixer.
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. Additional embodiments of the TAPS mixer are depicted in U.S. patent applications having the following Ser. Nos. 11/188,596; 11/188,470; 11/188,598; and, 11/188,595. 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.
It will be seen in the aforementioned patent applications that air is provided between the pilot mixer and a centerbody to cool an aft portion of the pilot mixer housing. Air is provided separately from an air manifold in a centerbody positioned between the pilot and main mixers to assist fuel penetration into the main mixer. It has been found that the separate air circuits for the pilot mixer and the main mixer unnecessarily occupy precious space in the mixer assembly. Further, the aft end of the centerbody adjacent the main mixer is subject to overheating and possible damage.
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 also desired that a single air manifold be utilized for providing cooling to both the main mixer and the pilot mixer. Besides enlarging the area available for the pilot mixer and its components, it is highly desirable that the air manifold provide cooling to the aft end of the main mixer.
In a first exemplary embodiment of the invention, a mixer assembly for use in a combustion chamber of a gas turbine engine includes a pilot mixer, a main mixer and a centerbody 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 centerbody includes a fuel manifold therein for providing fuel to the fuel injection ports and an air manifold therein for providing air to an aft portion of the pilot mixer.
In a second exemplary embodiment of the invention, a mixer assembly for use in a combustion chamber of a gas turbine engine includes a pilot mixer, a main mixer and a centerbody 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 centerbody includes a fuel manifold therein for providing fuel to the fuel injection ports and an air manifold therein for providing air to an aft portion of the annular cavity in the main mixer.
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 “High Pressure Gas Turbine Engine Having Reduced Emissions,” having Ser. No. 11/188,483 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 outer radial wall 138 of centerbody 106. It will be seen that outer radial wall 138 preferably also includes a ramp portion 142 located at a middle 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 outer radial wall 138 is preferably located within an axial length 145 of swirler housing 144.
It will be seen in
Swirled air may also be provided at upstream end 127 of annular cavity 126 via a series of passages formed in upstream wall 132 of swirler housing, as shown and described in a patent application entitled, “Mixer Assembly For Combustor Of A Gas Turbine Engine Having A Main Mixer With Improved Fuel Penetration, having Ser. No. 11/188,598, which is owned by the assignee of the present invention and hereby incorporated by reference. Rather, it is seen from
It will be understood that air flowing through first swirler 144 will be swirled in a first direction and air flowing through second swirler 146 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 swirlers 144 and 146, 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 the vanes in swirlers 144 and 146 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 height 151 of first swirler vanes 150 is preferably greater than height 160 of second swirler vanes 152. Accordingly, a relatively greater amount of air flows through first swirler 144 than through second swirler 146 due to the greater passage area therefor. The relative heights of swirlers 144 and 146 may be varied as desired to alter the distribution of air therethrough, so the sizes depicted are only illustrative.
Centerbody 106, as stated above, is located between pilot mixer 102 and main mixer 104 and preferably includes a fuel manifold 140 in flow communication with a fuel supply and an air manifold 141 in flow communication with air from high pressure compressor discharge outlet 69. In particular, an interior wall 143 of centerbody 106 defines air manifold 141 and is configured so that air is supplied to an aft portion of pilot mixer 102 and/or main mixer 104. In particular, it will be seen that a plurality of circumferentially spaced first passages 157 are provided in an aft portion 156 of interior wall 143 so that cooling air is provided from air manifold 141 therethrough to an aft flange 159 of pilot mixer housing 108. It will be appreciated that first passages 157 are preferably oriented substantially parallel to centerline axis 120 through mixer assembly 100.
In addition, it will be seen that a plurality of circumferentially spaced second passages 161 are preferably provided in aft portion 156 of interior wall 143 so that air is provided to assist in the separation of the fuel-air mixture from main mixer 104 and the fuel-air mixture from pilot mixer 102. Second passages 161 are preferably oriented at an axial angle to centerline axis 120 in a range of approximately 30-60°.
It will also be seen that air manifold 141 preferably provides air to a plurality of third passages 163 located in an outer radial portion 165 of centerbody 106. In this way, air flowing through air manifold 141 is provided to an aft portion of annular cavity 126 in main mixer 104. Third passages 163 are preferably spaced and oriented so as to have a predetermined multihole cooling arrangement. More specifically, it will be appreciated that third passages 163 will preferably be oriented at an axial angle to centerline axis 120 in a range of approximately 15-40° and at a tangential angle to centerline axis 120 in a range of approximately 0-80°. Accordingly, air provided to annular cavity 126 via third passages 163 is preferably swirled. The direction of swirl may be in the same direction as the air swirled by swirler arrangement 130 or in the opposite direction thereof.
In order to prevent outer radial portion 165 of centerbody 106 from overheating, it is preferred that some type of heat transfer item 167 be incorporated therewith. This can be in the form of ribs which extend circumferentially around an inner surface of outer radial portion 165, ribs which extend along an axial length of outer radial portion 165, or even a series of raised areas thereon having any desirable configuration.
It will further be noted that interior wall 143 serves to separate air manifold 143 and fuel manifold 140. An air gap 170 is still preferably provided between fuel manifold 140 and pilot housing 108. Fuel injection ports 128 are in flow communication with fuel manifold 140 and spaced circumferentially around centerbody 106. Contrary to that shown and described in a patent application entitled “Mixer Assembly For Combustor Of A Gas Turbine Engine Having A Main Mixer With Improved Fuel Penetration,” having Ser. No. 11/188,598 and also owned by the assignee of the present invention, fuel injection ports 128 are preferably positioned axially downstream of ramp portion 142 of centerbody outer radial wall 138. Regardless of the axial location of fuel injection ports 128, it is intended that the fuel be injected at least a specified distance into a middle radial portion of annular cavity 126 and away from the surface of inner wall 138.
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 outer 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 preferably 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.
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 air manifold 141. 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 the purge holes.
As shown in
When fuel is provided to main mixer 104, an annular, secondary combustion zone 198 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.
Contrary to the mixer assembly 100 depicted in the aforementioned '598, '470 and '596 patent applications, the configuration of air manifold 141 within centerbody 106 serves to eliminate the need for a separate air circuit previously provided immediately adjacent pilot housing 108. More specifically, it will be seen that air manifold 141, which is defined by interior wall 143, has been enlarged so as to provide air through first passages 157 to an aft flange 159 of pilot housing 108, through second passages 161 to an area between pilot mixer 102 and main mixer 104, through third passages 163 to annular cavity 126 of main mixer, and to inner passages 191 of each post member 190. In this way, the area available for pilot mixer 102 has been enlarged.
Regarding the relative distribution of air from air manifold 141, it will be understood that at least approximately half of such air supplied thereto is provided to inner passages 191 surrounding post members 190, and preferably at least 60% thereof is provided thereto. Accordingly, it is preferred that approximately 10-25% of the air supplied to air manifold 141 be provided to first passages 157, approximately 10-25% of the air supplied to air manifold 141 be provided to second passages 161, and approximately 10-25% of the air supplied to air manifold 141 be provided to third passages 163.
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.
