Patent Application
20070028595
The present invention relates to gas turbine engines having a high pressure ratio and, more particularly, to a staged combustion system for the gas turbine engine which is configured to minimize the production of undesirable combustion product components over the engine operating regime.
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.
While the design of the mixer assembly is able to improve mixing of fuel and air, and therefore reduce the emissions generated by the gas turbine engine, it has been found that the configuration and operation of the overall combustion system needs to be reconsidered if emissions are to meet desired levels without adversely affecting performance. This not only involves sizing the combustor properly, but also orienting and shaping the combustion chamber with respect to the mixer assemblies and the turbine nozzle. Further, the various hardware components of the combustor should be consistent with the air distribution requirements for cooling and lean burning, given the amount of compressed air flow provided to the combustor.
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, it is desired that such combustor retain required performance levels and characteristics. Further, 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. Modification of the combustor liners and combustion chamber is also desired so as to enable optimal use of the compressed air to the combustor.
In a first exemplary embodiment of the invention, a gas turbine engine having a longitudinal centerline axis therethrough is disclosed as including: a fan section at a forward end of the gas turbine engine for producing a first compressed air flow; a first compressor positioned downstream of the fan section and in flow communication with at least a portion of the first compressed air flow, wherein the first compressor produces a second compressed air flow having a designated pressure; a combustor positioned downstream of the first compressor and in flow communication with second compressed air flow, wherein the combustor produces combustion products from a mixture of fuel and air, a first turbine positioned downstream of the combustor and in flow communication with combustion products, wherein the first turbine powers the first compressor by means of a first rotatable drive shaft connected therebetween; and, a second turbine positioned downstream of the first turbine and in flow communication with the combustion products exiting the first turbine, wherein the second turbine powers the fan section by means of a second drive shaft connected therebetween. The gas turbine engine produces no more than a predetermined amount of emissions during an operating cycle.
In a second exemplary embodiment of the invention, a combustor of a gas turbine engine is disclosed as including: an annular dome portion at an upstream end having an outer end, an inner end and a plurality of circumferentially spaced openings therethrough; an outer liner connected to the outer end of the dome portion; an inner liner connected to the inner end of the dome portion and radially spaced from the outer liner to define a combustion chamber therebetween; a mixing assembly aligned with and located adjacent to each dome portion opening, and, a turbine nozzle located at a downstream end of the combustion chamber. The combustion chamber is configured so that a centerline axis through each mixing assembly is in substantial alignment with a center point of the turbine nozzle.
In accordance with a third embodiment of the present invention, a combustor for a gas turbine engine is disclosed as including: an annular dome portion at an upstream end having an outer end, an inner end and a plurality of circumferentially spaced openings therethrough; an outer liner connected to the outer end of the dome portion; an inner liner connected to the inner end of the dome portion and radially spaced from said outer liner to define a combustion chamber therebetween; a mixing assembly aligned with and located adjacent to each dome portion opening; and, a turbine nozzle located at a downstream end of the combustion chamber. The outer and inner liners only have openings therethrough in flow communication with the compressed air for cooling.
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 support struts 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 seen in
Low pressure turbine 32 preferably drives booster compressor 24 and fan rotor 38 via second drive shaft 34.
More specifically, combustion chamber 62 is housed within an engine outer casing 18 and is defined by an annular combustor outer liner 76, a radially-inwardly positioned annular combustor inner liner 78, and a dome plate 80 at its upstream end.
The arrows in
As seen in
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. Such mixing assemblies 67 are consistent with the TAPS mixers shown and described in the U.S. patents identified hereinabove. It will be seen that 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.
As seen in
More specifically, combustion chamber 162 is housed within an engine outer casing 118 and is defined by an annular combustor outer liner 176, a radially-inwardly positioned annular combustor inner liner 178, and a dome plate 180 at its upstream end. The arrows in
Contrary to the prior combustor, however, combustion chamber 162 thereof is generally symmetrical when viewed in cross-section and has a relatively larger dome height 163. This stems from the higher pressure ratios of the current gas turbine engines, which now are 30 and above. It will be appreciated by those skilled in the art that the pressure ratio of a gas turbine engine is generally defined as the ratio of second compressed air flow 56 (i.e., compressor discharge air) to an ambient pressure outside gas turbine engine 10. In fact, some gas turbine engines have pressure ratios greater than 40. It has been found that combustion chamber 162 should be sized according to the pressure ratio of gas turbine engine 10, where its volume and dome height 163 is increased as the pressure ratio of the gas turbine engine increases.
It is also preferred that combustion chamber 162, as well as outer liner 176, inner liner 178 and dome plate 180, be configured so that mixing assemblies 167 provided at the upstream end thereof have a centerline axis 169 therethrough in substantial alignment with a center portion of turbine nozzle vanes 174.
Outer and inner liners 176 and 178 of combustor 160 preferably are constructed so as to not include any dilution holes or openings. As seen in
In particular, the percentage of air provided to mixers 167 of the compressor discharge air supplied to combustor 160 is preferably greater than approximately 50-70% thereof. Approximately 4-6 times more air is provided to main mixer 104 (55-65% of compressor discharge air) than to pilot mixer 102 (8-15% of compressor discharge air). Thus, no more than approximately 30-40% of the compressor discharge air is provided as total cooling air for combustor 160. Of the total cooling air for combustor 160, approximately 5-15% of the compressor discharge air is provided as cooling air for dome 180 and approximately 15-25% thereof is provided as cooling air for outer and inner liners 176 and 178. It will also be appreciated that improvements in the variation in temperature along a given axial plane through outer and inner liners 176 and 178 have been experienced (no greater than approximately 140° F.). In addition, a pattern factor at a downstream end of combustion chamber 162 has improved to be no greater than approximately 1.1-1.3.
It will be seen in
With respect to mixers 167 in combustor 160, it is preferred that they have one of or a combination of the configurations and/or features shown and described in a group of patent applications filed concurrently herewith having the following titles: “Mixer Assembly For Combustion Chamber Of A Gas Turbine Engine Having A Plurality Of Counter-Rotating Swirlers,” having Ser. No. ______/______,______; “Swirler Arrangement For Mixer Assembly Of A Gas Turbine Engine Combustor Having Shaped Passages,” having Ser. No. ______/______,______; “Mixer Assembly For Combustor Of A Gas Turbine Engine Having A Main Mixer With Improved Fuel Penetration,” having Ser. No. ______/______,______; and, “Air-Assisted Fuel Injector For Mixer Assembly Of A Gas Turbine Engine Combustor,” having Ser. No. ______/______,______. Each of these applications is owned by the assignee of the present invention and are hereby incorporated by reference.
As seen 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 length 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
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, which is filed concurrently herewith and is owned by the assignee of the present invention. 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 188 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 188 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 length 151 of first swirler vanes 150 is preferably greater than length 180 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 lengths of swirlers 144 and 146 may be varied as desired to alter the distribution of air therethrough, so the sizes depicted are only illustrative.
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 200 of fuel manifold 106, and a shroud member 202 is configured to provide an inner surface 204 and an aft surface 206. Fuel injection ports 128 are in flow communication with fuel manifold and spaced circumferentially around centerbody outer shell 140. 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,” filed concurrently herewith and also owned by the assignee of the present invention, 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 188 before entering downstream end 131 of annular cavity 126. 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 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 185 is provided adjacent to and between each fuel injection port 128 to assist the injected fuel in its intended path. Such purge holes 185 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 210 having an inner passage 211 be associated with each such fuel injection port 128. It will be seen that post member 210 preferably extends from fuel injection port 128 through an air cavity 212 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 185.
As shown in
In light of the improvements made in combustor 160, gas turbine engine 10 produces no more than a predetermined amount of emissions during an operating cycle. More specifically, gas turbine engine produces no more than approximately 15-30 grams of NOx per kilogram of fuel and no more than approximately 5-10 grams of CO per kilogram of fuel during the take-off and landing portions of the operating cycle. It has also been found that gas turbine engine 10 produces no more than approximately 8-12 grams of NOx per kilogram of fuel during a cruise portion of the operating cycle. Further, no more than approximately 50-60 grams of unburned hydrocarbons per kilogram of fuel is produced during the ground idle portion of the operating cycle. Gas turbine engine has a smoke number of no more than approximately 1-10 during the take-off and landing portions of the operating cycle and a smoke number of no more than approximately 1-7 during the cruise portion of such operating cycle.
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.
