JP2009041848A - Pilot mixer for mixer assembly of gas turbine engine combustor including primary fuel injector and a plurality of secondary fuel injection ports - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To reduce discharge of harmful combustion components in a wide operating state of a gas turbine. <P>SOLUTION: The pilot mixer 102 includes: an annular pilot housing 108 having a hollow interior; the primary fuel injector 110 mounted inside the pilot housing 108 for supplying fuel droplets to the hollow interior of the pilot housing 108; a plurality of axial swirlers 112, 114 disposed upstream of the primary fuel injector 110; and the plurality of secondary fuel injection ports 134 for introducing fuel into the hollow interior of the pilot housing 108. A main mixer 104 further includes: a main housing 124 surrounding the pilot housing 108 and forming an annular hollow 126; a plurality of fuel injection ports 128 for introducing fuel into the hollow 126; and at least one swirler 130 disposed upstream of the plurality of fuel injection ports 128. <P>COPYRIGHT: (C)2009,JPO&INPIT

Description

  The present invention relates to a multi-stage combustion system in which the generation of harmful combustion product components is minimized over the entire operating region of the engine, and more specifically, has a pilot mixer with a primary fuel injector and a secondary fuel injection port. The present invention relates to a mixer assembly.

  Recent emphasis on minimizing the production and release of gases that cause smog and other hazardous environmental conditions, particularly those exhausted from gas turbine engines, is the production of such harmful combustion product components. And resulted in various combustor designs developed in an effort to reduce emissions. Another factor that impacts combustor design is the desire of gas turbine engine users for efficient and low cost operation, which reduces fuel consumption while maintaining or even increasing engine power Replaced by the need to do. As a result, important design criteria in aircraft gas turbine engine combustion systems include high thermal efficiency under various engine operating conditions, and also with precursors for particulate emissions, hazardous gas emissions, and photochemical smog formation. In order to obtain a minimization of the harmful combustion conditions that cause the emission of certain combustion products, the realization of a high combustion temperature is included.

  Various government regulatory agencies set emission limits for permissible levels of unburned hydrocarbons (HC), carbon monoxide (CO), and nitrogen oxides (NOx) that have been identified as the main factors in the generation of hazardous atmospheric conditions. ing. Accordingly, various combustor designs have been developed to meet these standards. For example, one way to address the problem of minimizing the emission of harmful combustion products in a gas turbine engine is to equip multi-stage combustion. Within this device, a combustor is provided that more precisely controls the characteristics of the combustion products using a first stage burner for low speed, low power conditions. For higher power conditions, a combination of first and second stage burners is provided, trying to keep the combustion products within emission limits. Balancing the operation of the first and second stage burners to enable efficient thermal operation of the engine while at the same time minimizing the production of harmful combustion products proves difficult to achieve. Will. In this regard, operating at low combustion temperatures to reduce NOx emissions can also result in incomplete or partially incomplete combustion, thereby producing excessive amounts of HC and CO, In addition, the output and thermal efficiency may be reduced. On the other hand, high combustion temperatures improve thermal efficiency and reduce the amount of HC and CO, but in many cases will further increase NOx emissions.

  Another method that has been proposed to minimize the production of these harmful combustion product components is to more effectively mix the injected fuel and combustion air. In this regard, numerous mixer designs that improve the mixing of fuel and air have been proposed over the years. In this way, combustion occurs uniformly throughout the mixture, reducing the level of HC and CO produced by incomplete combustion. However, even if mixing is improved, high levels of harmful NOx are formed at high power conditions if the flame temperature is high.

One mixer design that has been utilized is known as a twin annular premixed swirler (TAPS) and is described in U.S. Pat. Nos. 6,354,072, 6,363,726, 6,367. No. 6,262, No. 6,381,964, No. 6,389,815, No. 6,418,726, No. 6,453,660, No. 6,484,489 and No. 6,865,889 Is disclosed. US 2002/0178732 also shows some embodiments of a TAPS mixer. It will be appreciated that the TAPS mixer assembly includes a pilot mixer that is fueled during the entire engine operating cycle and a main mixer that is fueled only during the high power state of the engine operating cycle. U.S. Patents with improvements in the main mixer of the assembly during high power conditions (i.e. takeoff and ascent) have application numbers 11 / 188,596, 11 / 188,598 and 11 / 188,470 Although disclosed in the application, it is desirable to improve the pilot mixer to improve operability over other parts of the engine's operating envelope (i.e., idle, approach and cruise) while maintaining combustion efficiency. ing.
US Pat. No. 6,354,072 US Pat. No. 6,363,726 US Pat. No. 6,367,262 US Pat. No. 6,381,964 US Pat. No. 6,389,815 US Pat. No. 6,418,726 US Pat. No. 6,453,660 US Pat. No. 6,484,489 US Pat. No. 6,865,889 US Patent Application Publication No. 2002/0178732 US Patent Application Publication No. 11 / 188,596 US Patent Application Publication No. 11 / 188,598 US Patent Application Publication No. 11 / 188,470 US Patent Application Publication No. 11 / 188,483 US Patent Application Publication No. 11 / 188,595 US Pat. No. 6,405,523

  Accordingly, there is a need to provide a gas turbine engine combustor that minimizes the generation of harmful combustion product components over a wide range of engine operating conditions. Further, it is desirable that the pilot mixer of the nested combustor device be improved to improve operability and reduce emissions throughout the engine operating range.

  In a first exemplary embodiment of the present 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 to supply fuel droplets to the hollow interior of the pilot housing, A plurality of axial swirlers disposed upstream of the fuel injector, each of the plurality of swirlers swirling the air moving through the respective swirler and fuel supplied by the air and the primary fuel injector A plurality of axial swirlers having a plurality of vanes for mixing the droplets 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 forming an annular cavity, a plurality of fuel injection ports for introducing fuel into the cavity, and at least one swirler disposed upstream of the plurality of fuel injection ports. Each of the main mixer swirlers has a plurality of vanes for swirling the air moving through the respective swirler to mix the air and the fuel droplets supplied by the main mixer fuel injection port And at least one swirler. The fuel manifold is in flow communication with a plurality of secondary fuel injection ports in the pilot mixer and a plurality of fuel injection ports in the main mixer.

  In a second exemplary embodiment of the present invention, a method for operating a gas turbine engine combustor having a pilot mixer and a main mixer is disclosed, wherein the pilot mixer is an annular pilot housing having a hollow interior. A primary fuel injector mounted in the pilot housing and configured to supply fuel droplets into the hollow interior of the pilot housing, and a plurality of axial swirlers disposed upstream of the primary fuel injector. A plurality of axial directions having a plurality of vanes for swirling the air each moving through the respective swirler to mix the air and the fuel droplets supplied by the primary fuel injector A swirler and a plurality of secondary fuel injection ports for introducing fuel into the hollow interior of the pilot housing. The method includes supplying air at a specified flow rate through a swirler, supplying fuel via a primary fuel injector, and a secondary of the pilot mixer during a predetermined point in the gas turbine engine operating cycle. Supplying fuel via a fuel injection port.

  In a third exemplary embodiment of the present invention, a combustor for a gas turbine engine includes an outer liner and a combustion chamber spaced radially from the outer liner and between the outer liner. An inner liner adapted to form, a dome disposed at the upstream end of the combustion chamber, and a plurality of mixer assemblies disposed within the opening of the dome are disclosed. Each mixer assembly includes a pilot mixer, the pilot mixer being an annular pilot housing having a hollow interior, and a primary mounted in the pilot housing and adapted to supply fuel droplets into the pilot housing hollow interior. A fuel injector and a plurality of axial swirlers disposed upstream of the primary fuel injector, each of the swirlers swirling the air moving through the respective swirler to inject the air and the primary fuel injector A plurality of axial swirlers having a plurality of vanes for mixing the fuel droplets supplied by the vessel and a plurality of secondary fuel injection ports for introducing fuel into the hollow interior of the pilot housing.

  Referring now in detail to the drawings in which like reference numerals represent like elements throughout the drawings, FIG. 1 illustrates in schematic form an exemplary gas turbine engine 10 (high bypass type) for use in an aircraft. The gas turbine engine 10 has a longitudinal or axial central axis 12 therethrough for reference. The engine 10 preferably includes a core gas turbine engine identified generally by reference numeral 14 and a fan section 16 disposed upstream of the core gas turbine engine 14. The core engine 14 generally includes a generally tubular outer casing 18 that defines an annular inlet 20. The outer casing 18 further surrounds and supports a booster compressor 22 for raising the pressure of the air entering the core engine 14 to a first pressure level. A high-pressure multistage axial compressor 24 receives pressurized air from the booster 22 and further increases the pressure of the air. The pressurized air flows to the combustor 26 where fuel is injected into the pressurized air stream to increase the temperature and energy level of the pressurized air. The high energy combustion products flow from the combustor 26 to the first (high pressure) turbine 28 to drive the high pressure compressor 24 via the first (high pressure) drive shaft 30 and then the second (low pressure). ) Flows to the turbine 32 and drives the booster compressor 22 and the fan section 16 via a second (low pressure) drive shaft 34 coaxial with the first drive shaft 30. After driving each of the turbines 28 and 32, the combustion products exit the core engine 14 through the exhaust nozzle 36 and provide propulsion jet thrust.

  The fan section 16 includes a rotating axial fan rotor 38 surrounded by an annular fan casing 40. It will be appreciated that the fan casing 40 is supported by the core engine 14 by a plurality of generally radially extending and circumferentially spaced outlet guide vanes 42. In this way, the fan casing 40 surrounds the fan rotor 38 and the fan rotor blade 44. The downstream section 46 of the fan casing 40 extends above the outer portion of the core engine 14 to form a secondary or bypass airflow conduit 48 that provides additional propulsion jet thrust.

  From the flow point of view, the initial air flow represented by the arrow 50 flows into the gas turbine engine 10 via the inlet 52 to the fan casing 40. The air stream 50 flows through the fan blades 44 and travels through the conduit 48, a first pressurized air stream (represented by arrows 54), and a second pressurized air stream entering the booster compressor 22 ( (Represented by arrow 56). The pressure of the second pressurized air stream 56 is increased and enters the high pressure compressor 24 as represented by arrow 58. After being mixed with fuel and combusted in the combustor 26, the combustion product 60 exits the combustor 26 and flows through the first turbine 28. Combustion product 60 then flows through second turbine 32 and out of exhaust nozzle 36 to provide thrust for gas turbine engine 10.

  As best seen in FIG. 2, the combustor 26 includes an annular combustion chamber 62 coaxial with the longitudinal axis 12 and an inlet 64 and outlet 66. As pointed out above, the combustor 26 receives an annular pressurized air stream from the high pressure compressor outlet 69. A portion of this compressor discharge air flows into the mixing (mixer) assembly 67 where fuel is injected from the fuel nozzle 68 and mixed with the air to form a fuel-air mixture. The air-fuel mixture is supplied to the combustion chamber 62 for combustion. The fuel-air mixture is ignited by a suitable igniter 70, and the resulting combustion gas 60 flows axially toward the annular first stage turbine nozzle 72 and within the first stage turbine nozzle 72. Flowing into. The nozzle 72 is formed by an annular flow passage that includes a plurality of radially extending and circumferentially spaced nozzle vanes 74, which are arranged such that the gas flows at an angle and the first turbine 28. The gas is redirected to impinge on the first stage turbine blade. As shown in FIG. 1, the first turbine 28 preferably rotates the high pressure compressor 24 via a first drive shaft 30. The low pressure turbine 32 preferably drives the booster compressor 22 and the fan rotor 38 via the second drive shaft 34.

  The combustion chamber 62 is housed in the engine outer casing 18 and is formed by an annular combustor outer liner 76 and an annular combustor inner liner 78 disposed radially inward. The arrows in FIG. 2 indicate the direction in which the compressor discharge air flows in the combustor 26. As shown, a portion of the air flows over the outermost surface of the outer liner 76, a portion flows into the combustion chamber 62, and a portion flows over the innermost surface of the inner liner 78. .

  Unlike previous designs, the outer and inner liners 76, 78 each have additional air flowing into the combustion chamber 62 to complete the combustion process before the combustion products flow into the turbine nozzle 72. It is preferable not to provide a plurality of dilution ports that enable the above. This is in accordance with a patent application (Application No. 11 / 188,483) entitled “High Pressure Gas Turbine Engine with Reduced Emissions” owned by the applicant of the present invention. However, the outer liner 76 and the inner liner 78 are spaced in a plurality of smaller annulars that allow some of the air flowing along their outermost surfaces to flow into the interior of the combustion chamber 62. It will be appreciated that it is preferable to include a configured cooling air opening (not shown). These inwardly directed air flows will flow along the inner surfaces of the outer and inner liners 76 and 78 facing the interior of the combustion chamber 62 to form a film of cooling air along them.

It will be appreciated that a plurality of axially extending mixer assemblies 67 are arranged in a circular arrangement at the upstream end of the combustor 26 and extend into the inlet 64 of the annular combustion chamber 62. An annular dome plate 80 extends inwardly and forwardly to form the upstream end of the combustion chamber 62 and is spaced apart in a plurality of circumferentially formed therein to receive the mixer assembly 67. It will be seen that it has an aperture formed. For those portions, the upstream portion of each of the inner and outer liners 76 and 78, respectively,
Spaced radially apart from one another, an outer cowl 82 and an inner cowl 84 are formed. The spacing between the foremost ends of the outer and inner cowls 82 and 84 forms a combustion chamber inlet 64 and constitutes an opening that allows compressor discharge air to flow into the combustion chamber 62.

  FIG. 3 illustrates a mixing (mixer) assembly 100 according to one embodiment of the present invention. Mixer assembly 100 preferably includes a pilot mixer 102, a main mixer 104, and a fuel manifold 106 disposed between the two mixers. More specifically, the pilot mixer 102 includes an annular pilot housing 108 having a hollow interior and a primary fuel injector mounted within the housing 108 and adapted to supply fuel droplets into the hollow interior of the pilot housing 108. It will be appreciated that 110 is preferably included. Further, the pilot mixer 102 includes a first swirler 112 installed at a radially inner position adjacent to the primary fuel injector 110, and a second swirler 114 installed at a radially outer position of the first swirler 112. And a splitter 116 disposed between the swirlers. As shown, the splitter 116 extends downstream of the primary fuel injector 110 to form a venturi 118 in the downstream portion. The first and second pilot swirlers 112 and 114 are oriented substantially parallel to the central axis 120 of the mixer assembly 100 and include a plurality of vanes for swirling air moving through the swirlers 112 and 114. Will understand. The pilot mixer 102 is supplied with fuel and air at all times during the engine operating cycle so that a primary combustion zone 122 is created in the central portion of the combustion chamber 62 (see FIG. 2).

  The main mixer 104 further includes an annular main housing 124 that radially surrounds the pilot housing 108 and forms an annular cavity 126, a plurality of fuel injection ports 128 that introduce fuel into the annular cavity 126, and a reference numeral 130 as a whole. And the swirler device specified in 1. The swirler device 130 is a patent application entitled “Mixer assembly for a combustor of a gas turbine engine having a plurality of counter rotating swirlers”, both of which are owned by the present invention. 188, 596) and several patent applications (Application No. 11 / 188,595) entitled "Swirler Apparatus for Gas Turbine Engine Combustor Mixer Assembly with Molded Passage". It can be configured as any of the forms. However, it will be appreciated that in FIG. 3 the swirler device 130 preferably includes at least a first swirler 144 disposed upstream of the fuel injection port 128. As shown, the first swirler 144 is preferably oriented generally radially with respect to the central axis 120 of the mixer assembly 100. Note that the first swirler 144 includes a plurality of vanes 150 for swirling the air flowing between them. Since the vanes 150 are spaced substantially uniformly in the circumferential direction, a plurality of substantially uniform passages are formed between adjacent vanes 150. Further, it will be appreciated that the swirler 144 can include vanes having various configurations that cause the passage to be in the desired shape as disclosed in the above-identified '595 patent application.

  The swirler device 130 is also illustrated as including a second swirler 146 disposed upstream of the fuel injection port 128 and preferably oriented substantially parallel to the central axis 120. The second swirler 146 further includes a plurality of vanes 152 for swirling the air flowing between them. Although the vanes 152 are shown as being spaced approximately evenly in the circumferential direction, thereby forming a plurality of substantially uniform passages therebetween, such vanes 152 also show the passages. It can have various configurations to achieve the desired shape.

  As described above, the fuel manifold 106 is installed between the pilot mixer 102 and the main mixer 104 and is in flow communication with the fuel supply source. The fuel injection ports 128 are in flow communication with the fuel manifold 106 and are circumferentially spaced around the central body outer shell 140. As can be seen in FIG. 3, the fuel injection port 128 is preferably arranged so that fuel is supplied into the upstream end of the annular cavity 126.

  When fuel is supplied to the main mixer 104, an annular secondary combustion zone 198 is formed in the combustion chamber 62, which is spaced radially outward from the primary combustion zone 122 and the primary combustion zone 122. Are concentrically enclosed. Depending on the dimensions of the gas turbine engine 10, as many as twenty mixer assemblies 100 can be arranged in a circular array at the inlet 64 of the combustion chamber 62.

  As best seen in FIGS. 3, 4 and 6, the pilot mixer 102 further includes a plurality of spaced apart secondary fuel injection ports 134 that allow fuel to pass through the pilot housing. It is also introduced into the hollow interior 108. It will be appreciated that the secondary fuel injection ports 134 are preferably circumferentially spaced around the pilot housing 108 in a designated plane 136 that intersects the central axis 120 of the mixer assembly 100. Let's go. Although the plane 136 in which the secondary fuel injection port 134 is located is shown as being located in the flared portion 138 of the pilot housing 108 downstream of the splitter 116, the plane containing such secondary fuel injection port 134 is shown. It will be appreciated that can be located substantially at the downstream end of the splitter 116 or even upstream. In practice, the axial length of the splitter 116 can be changed so that its relationship to the position of the secondary fuel injection port 134 can be changed.

  Similarly, although the plane 136 is shown as being oriented substantially perpendicular to the central axis 120, the secondary fuel injection port 134 may be inclined so that the plane 136 is either upstream or downstream as desired. It can arrange so that it may become diagonal. Further, regardless of the axial position or orientation of the plane 136 containing the secondary fuel injection port 134, each such secondary fuel injection port 134 is individually oriented substantially perpendicular to the central axis 120, or It can be oriented at an acute angle in the upstream direction or at an obtuse angle in the downstream direction.

  Further, while the secondary fuel injection port 134 of the pilot mixer 102 is preferably in flow communication with the fuel manifold 106, it will be appreciated that the secondary fuel injection port 134 can receive fuel from a separate source. Will. As can be seen in FIG. 5, the secondary fuel injection port 134 can be incorporated into an integral fuel injection assembly 135 that includes the fuel injection port 128 of the main mixer 104. In any case, the fuel is typically secondary fuel injection port upon occurrence of a specific event (eg, at a specified operating cycle of the gas turbine engine 10, when the compressor discharge air 58 is at a specified temperature, etc.). 134 is injected into the hollow portion of the pilot housing 108. Depending on the requirements of the particular condition, the fuel may be a secondary fuel at a higher flow rate than the fuel injected via the primary fuel injector 110, at a lower flow rate, or at substantially the same flow rate as the fuel. It is injected through the injection port 134. Of course, this means that fuel is always supplied by the primary fuel injector 110, but it may be preferable to supply fuel to the pilot mixer 102 only through the secondary fuel injection port 134. It is assumed that.

  In this way, the pilot mixer 102 has greater flexibility during operation over low power conditions (ie idle, approach and cruise). Specifically, it can be seen that the pilot mixer 102 can generate power in the gas turbine engine 10 up to about 30% of maximum thrust when fuel is supplied only to the primary fuel injector 110. Let ’s go. In comparison, the pilot mixer 102 can generate power in the gas turbine engine 10 up to about 70% of the maximum thrust when fuel is similarly supplied to the secondary fuel injection port 134 as well.

  In order to facilitate the desired fuel spray into the hollow interior of the pilot housing 108, the passage 142 preferably surrounds each secondary fuel injection port 134 of the pilot mixer 102. Each passage 142 is in flow communication with pressurized air through a supply 154 adjacent to the fuel manifold 106. This air is supplied to allow injection of fuel spray into the pilot housing 108 instead of being forced along the inner surface 156 of the pilot housing 108. This is further enhanced by providing in each passage 142 a swirler 158 that provides a swirl for the air injected around the fuel spray.

  The vane 115 (see FIG. 6) of the outer pilot swirler 114 is also preferably configured such that air flowing therethrough is directed toward the inner surface 156 of the pilot housing 108. In this way, such air can interact better with the fuel supplied by the secondary fuel injection port 134. Accordingly, the vane 115 is preferably inclined at about 30 ° to about 60 ° with respect to the central axis 120. In this way, the flare angle 160 of the pilot housing 108 is close.

  When considering the addition of the secondary fuel injection port 134 into the pilot mixer 102, it will be appreciated that the air flow rate therethrough is preferably maintained at a ratio of about 10% to about 30%. . Further, such secondary fuel injection port 134 helps reduce emissions generated by mixer assembly 100 during operation of gas turbine engine 10. In particular, the combustor 26 can operate for a greater amount of time with only fuel supplied to the pilot mixer 102. It has also been found that it is desirable to supply more fuel to the radially outer position of the pilot mixer 102.

  It will be appreciated that in connection with the physical embodiment of the mixer assembly 100, a method of operating the combustor 26 having the pilot mixer 102 as described herein is also provided. More specifically, such a method includes the following steps: supplying air at a specified flow rate via pilot swirlers 112 and 114, and supplying fuel via primary fuel injector 110. Supplying fuel via a secondary fuel injection port 134 during a predetermined condition in the combustor 26 and / or during a predetermined operating cycle of the gas turbine engine 10. Further, such a method may include additional steps relating to the operation of the main mixer 104, i.e. supplying air via the main swirlers 144 and 146, predetermined conditions in the combustor 26 and / or gas turbine engine. Supplying fuel via a fuel injection port 128 during ten predetermined operating cycles. The fuel is generally supplied to the pilot mixer 102 via the secondary fuel injection port 134, and at the same time, the fuel is also supplied via the primary fuel injector 110, but the fuel is supplied to the secondary fuel injection port 134. It is also possible for certain conditions to be supplied only by, but not simultaneously by the primary fuel injector 110.

  While particular embodiments of the present invention have been illustrated and described, it would be obvious to those skilled in the art that various changes and modifications can be made without departing from the spirit of the invention. Accordingly, it is intended to embrace all such changes and modifications that fall within the scope of the invention.

1 is a schematic view of a high bypass turbofan gas turbine engine. 1 is a longitudinal cross-sectional view of a gas turbine engine combustor having a staged configuration. FIG. FIG. 3 is an enlarged cross-sectional view of the mixer assembly shown in FIG. 2. FIG. 4 is a rear perspective view of the mixer assembly shown in FIGS. 2 and 3. 5 is a rear perspective view of a portion of the mixer assembly shown in FIGS. FIG. 6 is a partial perspective view of the mixer assembly shown in FIGS. 2-4 taken along line 6-6 of FIG.

Explanation of symbols

10 Aircraft turbofan engine (the whole)
12 Longitudinal axis 14 Core gas turbine engine (all)
16 fan section 18 core engine outer casing 20 annular core engine inlet 22 booster compressor 24 high pressure compressor 26 combustor 28 first (high pressure) turbine 30 first (high pressure) drive shaft 32 second (low pressure) turbine 34 Second (low pressure) drive shaft 36 Exhaust nozzle 38 Fan rotor 40 Annular fan casing 42 Exit guide vane 44 Fan rotor blade 46 Downstream section of fan casing 48 Bypass air flow conduit 50 Arrow representing initial air flow 52 To fan casing Inlet 54 Arrow representing first (outside) pressurized air flow 56 Arrow representing second (inside) pressurized air flow 58 Arrow representing pressurized air flow to the high pressure compressor 60 Combustion product 62 Combustion chamber 64 Combustion chamber inlet 66 Combustion chamber outlet 67 Mixer assembly (all )
68 Fuel nozzle 69 High pressure compressor discharge port 70 Igniter 72 First stage turbine nozzle 74 Nozzle vane 76 Combustor outer liner 78 Combustor inner liner 80 Dome plate 82 Outer cowl 84 Inner cowl 100 Mixer assembly (all)
102 Pilot mixer (the whole)
104 Main mixer (the whole)
106 Fuel manifold 108 Pilot housing 110 Primary fuel injector of pilot mixer 112 First (inner) pilot swirler 114 Second (outer) pilot swirler 115 Outer pilot swirler vane 116 Splitter 118 Venturi 120 Central axis 122 of the mixer assembly Primary combustion zone 124 Main housing 126 Annular cavity 128 Fuel injection port 130 Swirl device (radial / axial)
134 Pilot Fuel Secondary Fuel Injection Port 135 Integrated Fuel Injection Assembly 136 Plane Containing Secondary Fuel Injection Port 138 Pilot Housing Flare 140 Central Body Outer Shell 142 Passage Surrounding Secondary Fuel Injection Port 144 First ( Radial) swirler 146 Second (axial) swirler 148 Annular passage 150 Radial swirler vane 152 Axial swirler vane 154 Air supply 156 Pilot housing inner surface 158 Swirler in passage 142 160 Pilot flare angle 198 Secondary combustion zone

Claims (10)

A mixer assembly (100) for use in a combustion chamber (62) of a gas turbine engine (10) comprising:
A pilot mixer (102), a main mixer (104), and a fuel manifold (106);
The pilot mixer (102)
An annular pilot housing (108) having a hollow interior;
A primary fuel injector (110) mounted within the pilot housing (108) and adapted to supply fuel droplets into a hollow interior of the pilot housing (108);
A plurality of axial swirlers (112, 114) disposed upstream of the primary fuel injector (110), each of the plurality of swirlers (112, 114) passing through a respective swirler (112, 114). A plurality of axial swirlers (112, 114) having a plurality of vanes for swirling the moving air to mix the air and the fuel droplets supplied by the primary fuel injector (110);
A plurality of secondary fuel injection ports (134) for introducing fuel into the hollow interior of the pilot housing (108);
The main mixer (104)
(1) a main housing (124) surrounding the pilot housing (108) and forming an annular cavity (126);
(2) a plurality of fuel injection ports (128) for introducing fuel into the cavity (126);
(3) At least one swirler (130) disposed upstream of the plurality of fuel injection ports (128), wherein each of the main mixer swirlers (130) swirls the air moving through the respective swirler At least one swirler (130) having a plurality of vanes for mixing the air and fuel droplets supplied by the main mixer fuel injection port (128),
The fuel manifold (106) is in flow communication with a plurality of secondary fuel injection ports (134) in the pilot mixer (102) and a plurality of fuel injection ports (128) in the main mixer (104). Yes,
Mixer assembly (100). The pilot mixer (102)
A first axial swirler (112) disposed upstream of the primary fuel injector (110) and having a plurality of vanes for swirling air traveling therethrough;
A second having a plurality of vanes disposed upstream of the primary fuel injector (110) and radially outward of the first axial swirler (112) for swirling air moving therethrough; An axial swirler (114) of
A splitter (116) disposed between the first and second axial swirlers (112, 114) and extending downstream to the venturi portion (118);
The mixer assembly (100) of claim 1.   The mixer assembly (100) of claim 2, wherein the second swirler (114) is configured to direct air toward an inner surface (156) of the pilot housing (108).   Secondary fuel injection ports (134) of the pilot mixer (102) are circumferentially spaced within a designated plane (136) that intersects the central axis (120) of the pilot housing (108). The mixer assembly (100) of any preceding claim.   The mixer assembly (100) of claim 1, wherein the secondary fuel injection port (134) of the pilot mixer (102) is oriented substantially perpendicular to a central axis (120) of the pilot housing (108). .   The mixer assembly (100) of claim 1, wherein a secondary fuel injection port (134) of the pilot mixer (102) is formed in a flared portion (138) of the pilot housing (108).   The pilot mixer (102) can generate power in the gas turbine engine (10) up to about 30% of maximum thrust when fuel is supplied only from the primary fuel injector (110). Item 12. The mixer assembly (100) of item 1.   The pilot mixer (102) generates power to the gas turbine engine (10) up to 70% of maximum thrust when fuel is supplied by the primary fuel injector (110) and secondary fuel injection port (134). The mixer assembly (100) of any preceding claim, wherein the mixer assembly (100) can be made.   The mixer assembly (100) of any preceding claim, further comprising a passageway (140) surrounding each secondary fuel injection port (134) of the pilot mixer (102) and in flow communication with an air supply.   The mixer assembly (100) of claim 1, wherein an air flow rate through the pilot mixer (102) is between 10% and 30%.

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