Patent Grant
7762073
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 mixer assembly having a pilot mixer with a primary fuel injector and secondary fuel injection ports.
Modem day emphasis on minimizing the production and discharge of gases that contribute to smog and to other undesirable environmental conditions, particularly those gases that are emitted from gas turbine engines, have led to different combustor designs that have been developed in an effort to reduce the production and discharge of such undesirable combustion product components. Other factors that influence combustor design are the desires of users of gas turbine engines for efficient, low cost operation, which translates into a need for reduced fuel consumption while at the same time maintaining or even increasing engine output. As a consequence, important design criteria for aircraft gas turbine engine combustion systems include provision for high combustion temperatures, in order to provide high thermal efficiency under a variety of engine operating conditions, as well as the minimization of undesirable combustion conditions that contribute to the emission of particulates, and to the emission of undesirable gases, and to the emission of combustion products that are precursors to the formation of photochemical smog.
Various governmental regulatory bodies have established emission limits for acceptable levels of unburned hydrocarbons (HC), carbon monoxide (CO), and oxides of nitrogen (NOx), which have been identified as the primary contributors to the generation of undesirable atmospheric conditions. Therefore, different combustor designs have been developed to meet those criteria. For example, one way in which the problem of minimizing the emission of undesirable gas turbine engine combustion products has been attacked is the provision of staged combustion. In that arrangement, a combustor is provided in which a first stage burner is utilized for low speed and low power conditions to more closely control the character of the combustion products. A combination of first stage and second stage burners is provided for higher power outlet conditions while attempting to maintain the combustion products within the emissions limits. It will be appreciated that balancing the operation of the first and second stage burners 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, in addition to producing lower power output and lower thermal efficiency. High combustion temperature, on the other hand, although improving thermal efficiency and lowering the amount of HC and CO, often results in a higher output of NOx
Another way that has been proposed to minimize the production of those undesirable combustion product components is to provide for more effective intermixing of the injected fuel and the combustion air. In that regard, numerous mixer designs have been proposed over the years to improve the mixing of the fuel and air. In this way, burning occurs uniformly over the entire mixture and reduces the level of HC and CO that result from incomplete combustion Even with improved mixing, however, higher levels of undesirable NOx are formed under high power conditions when the flame temperatures are high.
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. While improvements in the main mixer of the assembly during high power conditions (i.e., take-off and climb) are disclosed in patent applications have Ser. Nos. 11/188,596, 11/188,598, and 11/188,470, modification of the pilot mixer is desired to improve operability across other portions of the engine's operating envelope (i.e., idle, approach and cruise) while maintaining combustion efficiency.
Thus, there is a need to provide a gas turbine engine combustor in which the production of undesirable combustion product components is minimized over a wide range of engine operating conditions. Further, it is desired that the pilot mixer of a nested combustor arrangement be modified to improve operability and reduce emissions over the engine's operating envelope.
In a first exemplary embodiment of the invention, a mixer assembly for use in a combustion chamber of a gas turbine engine is disclosed as including a pilot mixer, a main mixer, and a fuel manifold. More specifically, the pilot mixer includes: an annular pilot housing having a hollow interior; a primary fuel injector mounted in the pilot housing and adapted for dispensing droplets of fuel to the hollow interior of the pilot housing; a plurality of anal swirlers positioned upstream from the primary fuel injector, each of the plurality of swirlers having a plurality of vanes for swirling air traveling through the respective swirler to mix air and the droplets of fuel dispensed by the primary fuel injector; and, a plurality of secondary fuel injection ports for introducing fuel into the hollow interior of the pilot housing. The main mixer further includes: a main housing surrounding the pilot housing and defining an annular cavity; a plurality of fuel injection ports for introducing fuel into the cavity; and, at least one swirler positioned upstream from the plurality of fuel injection ports, each of the main mixer swirlers having a plurality of vanes for swirling air traveling through the respective swirler to mix air and the droplets of fuel dispensed by the main mixer fuel injection ports. The fuel manifold is in flow communication with the plurality of secondary fuel injection ports in the pilot mixer and the plurality of fuel injection ports in the main mixer.
In a second exemplary embodiment of the invention, a method of operating a gas turbine engine combustor having a pilot mixer and a main mixer is disclosed, wherein said pilot mixer includes an annular pilot housing having a hollow interior, a primary fuel injector mounted in the pilot housing and adapted for dispensing droplets of fuel to the hollow interior of the pilot housing, a plurality of axial swirlers positioned upstream from the primary fuel injector, wherein each of the plurality of swirlers has a plurality of vanes for swirling air traveling through the respective swirler to mix air and the droplets of fuel dispensed by the primary fuel injector, and a plurality of secondary fuel injection ports for introducing fuel into the hollow interior of the pilot housing. The method includes the steps of providing air through the swirlers at a designated air flow rate, providing fuel through the primary fuel injector, and providing fuel through the secondary fuel injection ports of the pilot mixer during predetermined points in an operating cycle of the gas turbine engine.
In a third exemplary embodiment of the invention, a combustor for a gas turbine engine is disclosed as including an outer liner, an inner liner spaced radially from the outer liner so as to form a combustion chamber therebetween, a dome positioned at an upstream end of the combustion chamber, and a plurality of mixer assemblies positioned within openings of the dome. Each mixer assembly has a pilot mixer which includes: an annular pilot housing having a hollow interior; a primary fuel injector mounted in the pilot housing and adapted for dispensing droplets of fuel to the hollow interior of the pilot housing; a plurality of axial swirlers positioned upstream from the primary fuel injector, each of the plurality of swirlers having a plurality of vanes for swirling air traveling through the respective swirler to mix air and the droplets of fuel dispensed by the primary fuel injector; and, a plurality of secondary fuel injection ports for introducing fuel into the hollow interior of the pilot housing.
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” and having Ser. No. 11/188,483, 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. Swirler arrangement 130 may be configured in any of several ways, as seen in a patent application entitled “Mixer Assembly For Combustor Of A Gas Turbine Engine Having A Plurality Of Counter-Rotating Swirlers” having Ser. No. 11/188,596 and a patent application entitled “Swirler Arrangement For Mixer Assembly Of A Gas Turbine Engine Combustor Having Shaped Passages” having Ser. No. 11/188,595, both of which are assigned to the owner of the present invention. It will be seen in
Swirler arrangement 130 also is shown as including a second swirler 146 positioned upstream from fuel injection ports 128 and preferably oriented substantially parallel to centerline axis 120. Second swirler 146 further includes a plurality of vanes 152 for swirling the air flowing therebetween. Although vanes 152 are shown as being substantially uniformly spaced circumferentially, thereby defining a plurality of substantially uniform passages therebetween, such vanes 152 may also have different configurations so as to shape the passages in a desirably manner.
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. Fuel injection ports 128 are in flow communication with fuel manifold 106 and spaced circumferentially around centerbody outer shell 140. As seen 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.
As best seen in
Similarly, plane 136 is depicted as being oriented substantially perpendicular to centerline axis 120, but secondary fuel injection ports 134 may be positioned so that plane 136 is skewed so as to be angled either upstream or downstream as desired. Further, regardless of the axial position or orientation of plane 136 containing secondary fuel injection ports 134, each such secondary fuel injection port 134 may individually be oriented substantially perpendicular to centerline axis 120, oriented upstream at an acute angle, or oriented downstream at an obtuse angle.
It will further be seen that secondary fuel injection ports 134 of pilot mixer 102 preferably are in flow communication with fuel manifold 106, although it could receive fuel from a separate source. As seen in
In this way, pilot mixer 102 has greater flexibility during operation across the lower power conditions (i.e., idle, approach and cruise). In particular, it will be appreciated that pilot mixer 102 is able to power gas turbine engine 10 up to approximately 30% of maximum thrust when fuel is provided solely to primary fuel injector 110. By comparison, pilot mixer 102 is able to power gas turbine engine 10 up to approximately 70% of maximum thrust when fuel is provided to secondary fuel injection ports 134 as well.
In order to promote the desired fuel spray into the hollow interior of pilot housing 108, it is preferred that a passage 142 surround each secondary fuel injection port 134 of pilot mixer 102. Each passage 142 is in flow communication with compressed air via a supply 154 adjacent to fuel manifold 106. This air is provided to facilitate injection of the fuel spray into pilot housing 108 instead of being forced along an inner surface 156 thereof This may further be enhanced by providing a swirler 158 within each passage 142 which provides a swirl to the air injected around the fuel spray.
It is also preferred that vanes 115 of outer pilot swirler 114 (see
In this way, a flare angle 160 of pilot housing 108 is approximated.
Considering the addition of secondary fuel injection ports 134 in pilot mixer 102, it will be appreciated that the flow rate of air therethrough is preferably maintained at a rate of approximately 10% to approximately 30%. Further, such secondary injection ports 134 assist in reducing the emissions produced by mixer assembly 100 during the operation of gas turbine engine 10. In particular, combustor 26 is able to operate only with fuel being supplied to pilot mixer 102 for a greater time period Also, it has been found that providing more fuel at a radially outer location of pilot mixer 102 is desirable.
In conjunction with the physical embodiments of mixer assembly 100, it will be understood that a method of operating combustor 26 having pilot mixer 102 as described herein is also presented. More specifically, such method includes the following steps: providing air through pilot swirlers 112 and 114 at a designated flow rate; providing fuel through primary fuel injector 110; and, providing fuel through secondary fuel injection ports 134 during predetermined conditions in combustor 26 and/or an operating cycle of gas turbine engine 10. Further, such method may include additional steps with respect to the operation of main mixer 104, including: providing air through main swirlers 144 and 146; and, providing fuel through fuel injection ports 128 during predetermined conditions in combustor 26 and/or the operating cycle of gas turbine engine 10. While fuel will generally be provided to pilot mixer 102 through secondary fuel injection ports 134 when fuel is also being provided through primary fuel injector 110, there may be certain conditions when fuel is provided only by secondary fuel injection ports 134 and not concurrently by primary fuel injector 110.
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.
| Number | Name | Date | Kind |
|---|---|---|---|
| 5257502 | Napoli | Nov 1993 | A |
| 6354072 | Hura | Mar 2002 | B1 |
| 6363726 | Durbin et al. | Apr 2002 | B1 |
| 6367262 | Mongia et al. | Apr 2002 | B1 |
| 6381964 | Pritchard, Jr. et al. | May 2002 | B1 |
| 6389815 | Hura et al. | May 2002 | B1 |
| 6405523 | Foust et al. | Jun 2002 | B1 |
| 6418726 | Foust et al. | Jul 2002 | B1 |
| 6453660 | Johnson et al. | Sep 2002 | B1 |
| 6484489 | Foust et al. | Nov 2002 | B1 |
| 6711898 | Laing et al. | Mar 2004 | B2 |
| 6865889 | Mancini et al. | Mar 2005 | B2 |
| Number | Date | Country | |
|---|---|---|---|
| 20090113893 A1 | May 2009 | US |
