General Electric Patente

Gas turbine engine (10) nozzle and associated airfoil (123) designs are disclosed. An example airfoil (123) includes: a convex panel (920, 1120); a concave panel (1020); an outer band (930, 1030, 1130); and an inner band (940, 1040, 1140), the inner band (940, 1040, 1140) having an inner band (940, 1040, 1140) flow area (IBFA) and including a plurality of metering holes (910, 1010, 1110), the airfoil (123) characterized by the following equation, in which the IBFA, inner band (940, 1040, 1140) radius (IBR), exhaust gas temperature (EGT), and engine core speed (N), are related as follows: =IBFA1e−6m22IBR∗14∗N300HzEGT10002, wherein 0.0000002 meters squared (m²) ≤ IBFA ≤ 0.0000019 m², 0.213 m ≤ IBR ≤ 0.221 m, 988 degrees Celsius (C) ≤ EGT ≤ 1120 C, 306 Hertz (Hz) ≤ N ≤ 353 Hz, and 0.169 ≤ Equation 1 ≤ 16.643.

Gas turbine engine (10) nozzle and associated airfoil (123) designs are disclosed. An example airfoil (123) includes: a concave panel (1020); a convex panel (920, 1120); a baffle; an outer band (930, 1030, 1130); and an inner band (940, 1040, 1140), the concave panel (1020) including a plurality of metering holes (910, 1010, 1110), the airfoil (123) characterized by the following equation, in which a concave panel (1020) metering hole area (1230) (CCVA (830, 1080)), baffle area (BA), inner band (940, 1040, 1140) radius (IBR), and loading area (840, 1090) (LA) are related as follows: EQ 1 = CCVA m 2 − 0.0000152 m 2 0.0000152 m 2 − 3 + BA m 2 0.001 m 2 LA m 2 0.00001 m 2 * IBR m 20 m , <img class="EMIRef" id="1219ed77-300e-4dab-b90a-d7bf0a7ea93a-ia01" /> wherein 0.00005 meters squared (m<2>) ≤ BA ≤ 0.0002 m<2>, 0.000005 m<2> ≤ CCVA ≤ 0.000015 m<2>, 0.213 m ≤ IBR ≤ 0.221 m, 0.0000147 m<2> ≤ LA ≤ 0.0000153 m<2>, and 4850.690 ≤ Equation 1 ≤ -0.03.

Gas turbine engine (10) nozzle and associated airfoil (123) designs are disclosed. An example airfoil (123) of an engine nozzle segment (140, 800) includes: a convex panel (920, 1120); a concave panel (1020); an outer band (930, 1030, 1130); and an inner band (940, 1040, 1140). The convex panel (920, 1120) includes a plurality of metering holes (910, 1010, 1110), and the airfoil (123) is characterized by the following equation, in which a convex panel (920, 1120) metering hole area (1230) (CVXA (860, 980)), inner band (940, 1040, 1140) radius (IBR), exhaust gas temperature (EGT), and engine core speed (N), are related as follows: =280∗CVXAm25x10−6m2+5x10−6IBRm1m∗3.8∗EGT°C1000°C∗NHz300Hz2 wherein 0.0000001 meters squared (m²) ≤ CVXA (860, 980) ≤ 0.000004 m², 0.213 m ≤ IBR ≤ 0.221 m, 988 degrees Celsius (C) ≤ EGT ≤ 1120 C, 306 Hertz (Hz) ≤ N ≤ 353 Hz, and 4.835 ≤ Equation 1 ≤ 281.150.

Gas turbine engine (10) nozzle and associated airfoil (123) designs are disclosed. An example airfoil (123) of an engine nozzle segment (140, 800) includes: a concave panel (1020); a convex panel (920, 1120); a baffle; an outer band (930, 1030, 1130); and an inner band (940, 1040, 1140), the concave panel (1020) including a plurality of metering holes (910, 1010, 1110), the airfoil (123) characterized by the following equation, in which a concave panel (1020) metering hole area (1230) (CCVA (830, 1080)), baffle area (BA), inner band (940, 1040, 1140) radius (IBR), exhaust gas temperature (EGT), and engine core speed (N), are related as follows: =CCVAm2−1.52x10−5m21.52x10−5m2−3+BAm20.001m2∗IBRm20m∗NHz300HzEGT°C1000°C2, wherein 0.00005 meters squared (m²) ≤ BA ≤ 0.0002 m², 0.000005 m² ≤ CCVA ≤ 0.000015 m², 0.213 m ≤ IBR ≤ 0.221 m, 988 degrees Celsius (C) ≤ EGT ≤ 1120 C, 306 Hertz (Hz) ≤ N ≤ 353 Hz, and -6900.49 ≤ Equation 1 ≤ -0.03.

Gas turbine engine (10) nozzle and associated airfoil (123) designs are disclosed. An example airfoil (123) of an engine nozzle segment (140, 800) of a gas turbine engine (10) includes: a convex panel (920, 1120); a concave panel (1020); an outer band (930, 1030, 1130); and an inner band (940, 1040, 1140). The convex panel (920, 1120) and the concave panel (1020) include a plurality of metering holes (910, 1010, 1110).

Gas turbine engine (10) nozzle and associated airfoil (123) designs are disclosed. An example airfoil (123) includes: a convex panel (920, 1120); a concave panel (1020); an outer band (930, 1030, 1130); and an inner band (940, 1040, 1140), the concave panel (1020) having an area of maximum curvature (MC) including a plurality of metering holes (910, 1010, 1110), the airfoil (123) characterized by the following equation, in which a MC metering hole area (1230) (MCA (870, 1180)), inner band (940, 1040, 1140) radius (IBR), and loading area (840, 1090) (LA) are related as follows: EQ 1 = IBR m 1 m LA m 2 0.00001 m 2 ∗ MCA 870 1180 m 2 0.000005 m 2 2 , <img class="EMIRef" id="34a8ea57-9ab3-45ca-be72-92fda08b5d6b-ia01" /> wherein 1.720x10<-6> meters squared (m<2>) ≤ MCA ≤ 3.000x10<-6> m<2>, 0.213 m ≤ IBR ≤ 0.221 m, 0.0000147 m<2> ≤ LA ≤ 0.0000153 m<2>, and 0.351 ≤ Equation 1 ≤ 1.644.

Gas turbine engine (10) nozzle and associated airfoil (123) designs are disclosed. An airfoil (123) of an engine nozzle segment (140, 800) includes: a convex panel (920, 1120); a concave panel (1020); an outer band (930, 1030, 1130); and an inner band (940, 1040, 1140). The convex panel (920, 1120) includes a plurality of metering holes (910, 1010, 1110), the airfoil (123) characterized by the following equation, in which a convex panel (920, 1120) metering hole area (1230) (CVXA (860, 980)), inner band (940, 1040, 1140) radius (IBR), and loading area (840, 1090) (LA) are related as follows: EQ1=CVXAm20.000005m2+0.000005IBRmπmLAm20.00001m2, wherein 0.0000001 meters squared (m²) ≤ CVXA ≤ 0.000004 m², 0.213 m ≤ IBR ≤ 0.221 m, 0.0000147 m² ≤ LA ≤ 0.0000153 m², and 4.835 ≤ Equation 1 ≤ 281.150.

Gas turbine engine (10) nozzle and associated airfoil (123) designs are disclosed. An example airfoil (123) includes: a convex panel (920, 1120); a concave panel (1020); an outer band (930, 1030, 1130); and an inner band (940, 1040, 1140), the concave panel (1020) having an area of maximum curvature (MC) that includes a plurality of metering holes (910, 1010, 1110), the airfoil (123) characterized by the following equation, in which a MC metering hole area (1230) (MCA (870, 1180)), inner band (940, 1040, 1140) radius (IBR), exhaust gas temperature (EGT), and engine core speed (N), are related as follows: Equation1=MCAm20.000005m22IBRm1m∗14∗NHz300HzEGT°C1000°C2, wherein 1.720x10^-6 meters squared (m²) ≤ MCA ≤ 3.000x10^-6 m², 0.213 m ≤ IBR ≤ 0.221 m, 988 degrees Celsius (C) ≤ EGT ≤ 1120 C, 306 Hertz (Hz) ≤ N ≤ 353 Hz, and 0.351 ≤ Equation 1 ≤ 1.644.

A system (100) includes an apparatus (130) for building a component by additive manufacturing, a camera (136), an adjustable optical attenuator (144), and a computing device (156). When the computing device (156) receives a signal indicating that a layer of particulate has been deposited onto a build plane (112), the attenuator (144) adjusts the attenuation level of the camera (136) to a low attenuation level, and the camera (136) captures a pre-weld image. When the computing device (156) receives a signal indicating a start of exposure of a layer of the component (134) to an energy beam (132), the attenuator (144) adjusts the attenuation level of the camera (136) to a high attenuation level, and the camera (136) captures an in-weld image. When the computing device (156) receives a signal indicating an end of exposure of the layer of the component (134) to the energy beam (132), the attenuator (144) adjusts the attenuation level of the camera (136) to the low attenuation level, and the camera (136) captures a post-weld image.

An apparatus (156) may include one or more processors (154), and memory (158) comprising machine-readable instructions that, when executed by the one or more processors (154), cause the apparatus (156) to receive a plurality of images of a particulate (114) on a build plane (112) of an additive manufacturing system (100) captured by a camera (136) while one layer of the particulate (114) is exposed to an energy beam (132) of the additive manufacturing system (100), and combine the plurality of images to generate an image of a melt pool trajectory of the one layer of the particulate (114).

A fuel nozzle assembly (48) for a gas turbine engine (10) is provided, the fuel nozzle assembly (48) comprising a first shaft (84) including a first fluid passage (100); a flange (82) connected to the first shaft (84) and including a second fluid passage (102) fluidly coupled with the first fluid passage (100); a nozzle tube (70) connected to the flange (82) and including a third fluid passage (104) fluidly coupled with the second fluid passage (102), the nozzle tube (70) disposed at a first side of the flange (82); and a second shaft (86) connected to the flange (82), the first shaft (84) disposed at least partially in the second shaft (86). The flange (82), the first shaft (84), and the second shaft (86) at least partially define a leak path (LP). The leak path (LP) includes a radial gap (212) between the first shaft (84) and the second shaft (86); and a leak channel (154) in an axial surface (156) of at least one of the first shaft (84) or the second shaft (86).

A system (100') includes an apparatus (130) for building a component by additive manufacturing using an energy beam (132), at least one first camera (136) having a first field of view of the build plane (112'), at least one second camera (136') having a second field of view of the build plane (112'), and a computing device (156). When the computing device (156) receives a first image (1100) and a second image (1102) of the build plane (112') captured by the at least one first camera (136) and second camera (136'), the images (1100, 1102) include a fiducial marker (1106) positioned on the build plane (112') within an overlap region (1104). Coordinates of the fiducial marker (1106) in the first image (1100) within the overlap region (1104) and within the second image (1102) are then identified. The first image (1100) and the second image (1102) are combined to form a merged image (1108) based on the identified coordinates of the fiducial markers (1106) in the first and second images (1100, 1102).

A turbine engine (10) includes a compression section (12), a combustion section (14), and a turbine section (16) in serial flow arrangement. The combustion section (14) has a combustor (30) including a combustor liner (40) at least partially defining a combustion chamber (74) exhausting at a combustion chamber outlet (76), with the combustor liner (40) having an inner liner (42) and an outer liner (44). A liquid fuel injector (70) is configured to inject a first mixture comprised of liquid fuel and air or atomized liquid fuel and air to the combustion chamber (74) and a gaseous fuel injector (72) is configured to inject a second mixture comprised of gaseous fuel and air or atomized fuel and air to the combustion chamber (74).

A system (100) includes an apparatus (130) for building a component by additive manufacturing, a camera (136), and a computing device (156). When the computing device (156) receives an image (600) of the build plane (112) captured by the camera (136), the image (600) includes a fiducial marker (602) positioned on the build plane (112), the computing device (156) identifies the coordinates of the fiducial marker (602) in the image in a coordinate system of the camera (136). The computing device (156) then identifies corresponding coordinates of the fiducial marker (602) in the coordinate system of the apparatus (130), and generates an image transfer function that converts the coordinates in the camera system to coordinates of the apparatus. The image transfer function is then applied to the image (600) to generate a distortion-corrected image (604).

An apparatus for additively manufacturing three-dimensional objects includes one or more laser beam sources configured to generate a laser beam, and one or more modulation devices disposed downstream of the one or more laser beam sources and configured to receive the laser beam and modulate the laser beam to generate a modulated beam. One or more computing systems are configured to generate one or more output images based on an input to the one or more computing systems of a wavefront pattern, wherein the one or more output images are generated by at least one of a graphics processing unit (GPU) or an integrated circuit. The one or more computing systems are configured to control the one or more modulation devices to generate the modulated beam with the wavefront pattern based on at least one output image of the one or more output images.

A system (100) includes an apparatus (130) for building a component by additive manufacturing, a camera (136), and a computing device (156). When the computing device (156) receives an image (600) of the build plane (112) captured by the camera (136), the image (600) includes a fiducial marker (602) positioned on the build plane (112). The computing device (156) rectifies the image (600) to generate a distortion-corrected image (604), and identifies the coordinates of the fiducial marker (602) in the distortion-corrected image (604). The computing device (156) then compares identified coordinates of the at least one fiducial marker (602) from the distortion-corrected image (604) with corresponding nominal coordinates of the at least one fiducial marker (602) in a coordinate system of the apparatus (130), and generates a coordinate transfer function based on a result of the comparison. The computing device (156) then applies the coordinate transfer function to the apparatus (130) to calibrate a position of the energy beam (132) on the build plane (112).

An aircraft (10) includes an engine (20) defining a centerline axis (24), the engine (20) comprising a fan (22) and a turbomachine (54) rotatably driving the fan (22), the turbomachine (54) including an exhaust section (36) comprising an outlet nozzle (38), the outlet nozzle (38) including a fixed portion (40) and a movable portion (42), wherein the movable portion (42) is movable from a first position to a second position, wherein when the movable portion (42) is in the first position, the movable portion (42) is aligned with the centerline axis (24), and wherein when the movable portion (42) is in the second position, the movable portion (42) is canted downward in the vertical direction and outward in the lateral direction relative to the centerline axis (24), and wherein the movable portion (42) includes an end (44) engaging the fixed portion (40), wherein when the movable portion (42) is in the first position, the end (44) defines a nonzero cant angle with the centerline axis (24).

A bleed valve system (80, 200, 300) for a gas turbine engine (10, 100, 100', 210, 240, 270), the bleed valve system (80, 200, 300) including a bleed valve (202, 202', 82, 222, 252, 262, 274, 310), a bleed valve exhaust duct (204, 304) fluidly connected to the bleed valve (202, 202', 82, 222, 252, 262, 274, 310), and a bleed valve exhaust nozzle (206, 306) fluidly connected to the bleed valve exhaust duct (204, 304), wherein the bleed valve exhaust duct (204, 304) and the bleed valve exhaust nozzle (206, 306) are fluidly separated from an exhaust nozzle of the gas turbine engine (10, 100, 100', 210, 240, 270) when the bleed valve system (80, 200, 300) is installed in the gas turbine engine (10, 100, 100', 210, 240, 270).

A frame assembly includes a frame (170) and a mounting bracket assembly (300, 302, 304, 306, 500, 502). The frame (170) includes an inner hub (174), an outer shell (172, 470) located opposite the inner hub (174), and a plurality of struts (158) connecting the inner hub (174) with the outer shell (172, 470). At least one strut of the plurality of struts (158) is a hollow strut (400) with a radial passage (440) extending therethrough. The mounting bracket assembly (300, 302, 304, 306, 500, 502) includes an inner flange (330) located on an inner surface (178) of the inner hub (174), an outer flange (320) located on an outer surface (176) of the outer shell (172, 470), and a radial linkage (340) connecting the inner flange (330) with the outer flange (320). The radial linkage (340) extends through the radial passage (440) of the hollow strut (400).

A turbine engine (100) for an aircraft includes an integral frame. The integral frame includes an inner hub (240), a frame case (220) having an outer band portion (222) and a containment portion (224) integrally formed with the outer band portion (222), and a plurality of struts (210) connecting the inner hub with the outer band portion. The outer band is opposite the inner hub to form an air flow path (187) therebetween, and the containment portion extends in an axial direction of the turbine engine away from the outer band. The containment portion is positioned radially outward of a plurality of rotating airfoils (262) and extends axially over the plurality of rotating airfoils. The outer band portion and the containment portion of the frame case can be an integral composite having a plurality of reinforcing fibers (272) embedded in a matrix (270) or an integral metallic part.

A propulsion assembly includes a first torque source coupled with a first shaft and a second torque source coupled with a second shaft. A coupler selectively couples the first and second torque sources. When the first and second torque sources are coupled via the coupler, in response to a command to decouple the first torque source, an unloading operation is performed to decrease the torque output provided by the first torque source to a threshold, and when reached, the first shaft is decoupled from the coupler. When the first torque source is coupled with the coupler but the second torque source is not, in response to a command to couple the second torque source, a speed matching operation is performed to increase the speed of the second shaft to match a speed of the first shaft, and when the speeds are matched, the second shaft is coupled to the coupler.

A hybrid electric propulsion system (10, 50, 100, 150, 200, 250, 300) includes a gearbox (16, 58, 104, 204, 302), an electric motor (52, 103, 203, 253) operably coupled to the gearbox (16, 58, 104, 204, 302), a battery (101, 251) operably coupled to the electric motor (52, 103, 203, 253) to power the electric motor (52, 103, 203, 253), and an engine lubrication system. The engine lubrication system includes a tank (108) that holds an engine lubrication fluid. The engine lubrication system includes a pump (110, 312) that is configured to pump the lubrication fluid from the tank (108) toward one or both of the battery (101, 251) and the electric motor (52, 103, 203, 253) through a supply line to maintain one or both of the battery (101, 251) and the electric motor (52, 103, 203, 253) above a threshold temperature.

A method of manufacturing a composite component having an outer shell, an inner hub, and a plurality of struts connecting the outer shell and the inner hub. A plurality of outer shell preform portions having bifurcated strut portions are connected together to form an outer shell hoop preform, and a plurality of inner hub preform portions having bifurcated strut portions are connected together to form an inner hub hoop preform. The bifurcated strut portions are arranged to extend between the outer shell hoop preform and the inner hub hoop preform, and are arranged adjacent to one another to form a strut preform. A matrix material is injected into the mold tooling structure and a curing process is applied to the mold tooling structure to obtain the composite component.

A method of manufacturing a composite component (161') having an outer shell (165), an inner hub (167), and a plurality of struts (158) connecting the outer shell (165) and the inner hub (167). An outer shell frame preform (302, 308) is woven to include a plurality of outer shell strut preform portions (318), each including an outer shell strut leading edge preform portion (410) and an outer shell strut trailing edge preform portion (412). An inner hub frame preform (304, 312) is woven to include a plurality of inner hub strut preform portions (336), each including an inner hub strut leading edge preform portion (478) and an inner hub strut trailing edge preform portion (480). In forming a strut preform (306), the outer shell leading edge preform portion (410) and the inner hub leading edge preform portion (478) are connected together, and the outer shell trailing edge preform portion (412) and the inner hub trailing edge preform portion (480) are connected together.

An apparatus for additively manufacturing three-dimensional objects includes a laser beam source configured to generate a laser beam, and a modulation device disposed downstream from the laser beam source. The modulation device includes a dichroic layer and a modulation layer. The dichroic layer is configured to transmit a first wavelength range of the laser beam and a second wavelength range of the laser beam to the modulation layer and reject wavelengths of the laser beam falling outside the first and second wavelength ranges where the first wavelength range is a multiple of the second wavelength range. One or more optical devices are disposed downstream of the modulation device and configured to selectively scan the laser beam onto a build plane.

An additive manufacturing apparatus (10) includes a build module (22). A feed module (24) is configured to support a first portion of a resin support (28). The first portion of the resin support (28) is supported by a feed mounting panel (68). A take-up module (26) is configured to support a second portion of the resin support (28). The second portion of the resin support (28) is supported by a take-up mounting panel (74) and is positioned on an opposing side of the radiant energy device from the feed module (24). An adjustment assembly is configured to adjust a position of at least one of a feed mandrel (24A) within the feed module (24) or a take-up mandrel (26A) within the take-up module (26).

A system (100) for using large language models (LLMs) to convert natural language (NL) requirements for aerospace systems to formalism for controlling a landing gear system of an aircraft includes a processor (305) and a memory (330) including instructions which, when executed by the processor, causes the system at least to perform using an LLM to revise a received NL expression associated with a requirement for an aerospace system. A LLM is also used to replace at least one word in the revised NL expression with a lifted atomic entity to generate an underlying sentence structure and translate the underlying sentence structure to a lifted logical formula. A LLM is also used to ground the lifted logical formula and to convert the NL expression into a formalized requirement for the aerospace system. The system also controls the aerospace system based on the formalized requirement.

The present disclosure relates to systems and methods for improved anomaly detection for rotating machines. An example method may include determining a rotational speed of a rotating machine. The example method may also include determining, using a frequency domain transform of a signal of the rotating machine, a frequency domain signal. The example method may also include determining, based on the rotational speed of the rotating machine, a first frequency band within the frequency domain signal for identifying a fault frequency of the rotating machine. The example method may also include determining a fault frequency of the rotational machine within the first frequency band. The example method may also include determining, based on the fault frequency, a second frequency band within the first frequency band, wherein the second frequency band includes the fault frequency. The example method may also include determining, based on the second frequency band, a first fault index and a baseline of the first fault index. The example method may also include determining, based on a deviation of a second fault index from the baseline, a fault condition of the rotating machine. The example method may also include providing an alert based on the fault condition of the rotating machine.

A trained machine learning model identifies that a real-world apparatus has a failed component, which trained machine learning model has been trained with a training corpus that includes content generated by synthesizing a plurality of synthesized operating examples for a given apparatus, wherein at least some of the plurality of synthesized operating examples are generated via a simulation modeling environment that receives as input characterizing information that corresponds to any of a variety of failure states for a component of the given apparatus.

A gearbox assembly (46) includes a planet gear (106) including an outer surface (144) and an inner surface (140) opposite the outer surface (144). A pair of circumferential oil channels (136) are formed in the inner surface (140) proximate opposite ends of the planet gear (106). A distributor channel (138) is formed to extend in a longitudinal direction toward a center of the planet gear (106) from each circumferential oil channel (136). One or more tooth channels (142) are formed to extend from each distributor channel (138) to the outer surface (144) of the planet gear (106). A journal pin (112) is received within the planet gear (106) extending along an axis (13) about which the planet gear (106) rotates. A gap (132) defines a journal bearing (133) formed between the inner surface (140) of the planet gear (106) and the journal pin (112). An internal bore (114) is formed in the journal pin (112) extending from an end of the journal pin (112) to the gap (132).

A method for additively manufacturing three-dimensional objects includes modulating a first and second sets of modulation segments of a modulation device to respective different first and second modulation states. The first set of modulation segments is addressed via one or more laser beams. The one or more laser beams are switched from addressing the first set of modulation segments to addressing the second set of modulation segments via at least one of steering the one or more laser beams from the first set of modulation segments to the second set of modulation segments or switching from a first laser beam source generating the one or more laser beams addressing the first set of modulation segments to a second laser beam source generating the one or more laser beams addressing the second set of modulation segments. The one or more laser beams are directed onto a powder bed.

A turbine engine (10) includes an engine core (11) extending along an engine centerline (20) and includes a compressor section (12), a combustor (14), and a turbine section (16) in serial flow arrangement. A set of blades (30) are circumferentially arranged in the turbine section (16) and the compressor section (12). A set of dovetails (52) mounts the set of blades (30) to a disk (32), rotated about the engine centerline (20). Each dovetail (52) can include a first upper lobe (110) and a first lower lobe (112) defining a first intervening recess (114), and a second upper lobe (120) and a second lower lobe (122) defining a second intervening recess (124), collectively defining a neck (130) for mounting to the disk (32).

An insertion tool (105) for inspection is provided, including a tube (116) with a flexible section (118) having at least two joints (122) for articulating the flexible section (118) into different shapes, the flexible section (118) selectively configurable between an unrigidized state and at least a first rigid state having a first non-linear shape. The tool (105) includes an end effector (130) coupled to a distal end of the tube (116) with a sensor head (134), the sensor head (134) to direct a beam toward a target area of the equipment to perform a profilometry operation. The tool (105) includes a waveguide (128) extending through the tube (116), the waveguide to transmit the beam from an electromagnetic source (108) to the end effector (130) and to transmit electromagnetic signals from the end effector to a profilometry device, and an end effector actuator (112) to move the beam at the target area. The end effector (130) is axially spaced from an insertion axis of the tool (105) when the flexible section (118) is in the first rigid state to obtain off-axis profilometry measurements at the target area.

A turbine engine (10) including a combustor (34) with a combustor liner (38, 138) having dilution openings (62, 76, 78) and a geometry that changes along an axial direction. The combustor further having a baffle (70, 170, 270, 370) surrounding a combustor liner (38, 138) defining a combustion chamber (46) of the combustor. A method for controlling nitrogen oxides within the combustor, including injecting compressed air into the annular combustion chamber (46) through any of the dilution openings (62, 76, 78) described herein.

An engine (10) can utilize a combustor (36) to combust fuel to drive the engine (10). A fuel nozzle assembly (130) can supply fuel to the combustor (36) for combustion or ignition of the fuel. The fuel nozzle assembly (130) can include a swirler (134) and a fuel nozzle (132) to supply a mixture of fuel and air for combustion. The fuel nozzle assembly (130) can be configured to increase lateral provision of fuels to reduce flame scrubbing on combustor liners (40, 42) for the combustor (36).

A turbine engine (10) can combust fuel to drive a turbine, which drives the engine. A fuel nozzle (100), (200), (300), (400), (500) can supply fuel for combustion or ignition of the fuel. The fuel nozzle (100), (200), (300), (400), (500) can include a central body (110) defining a first fuel passage (111). An annular first wall (121) circumferentially surrounds the central body (110) and is spaced therefrom to define a first airflow passage (131) therebetween. A set of first swirler vanes (104) is circumferentially spaced about the central body (110) and disposed within the first airflow passage (131). The fuel nozzle (100), (200), (300), (400), (500) can supply a mixture of fuel and air for combustion.

A gas turbine engine (10), comprising: a compressor section (12), combustion section (14), and turbine section (16) in serial flow arrangement, with the combustion section (14) comprising: a combustor liner (40) at least partially defining a combustion chamber (50); a wall (46) coupled to the combustor liner (40); a first fuel supply (34) to supply a first fuel; a gaseous fuel supply (36) to supply a gaseous hydrogen fuel; and a fuel nozzle (100) assembly coupled to the wall (46) and fluidly coupled to the first fuel supply (34) and the gaseous fuel supply (36), the fuel nozzle (100) assembly comprising: a main mixer (102); and a fuel nozzle (100) disposed such that the main mixer (102) is disposed at least partially around the fuel nozzle (100).

A gas turbine engine (10), comprising a compressor section (12), combustion section (14), and turbine section (16) in serial flow arrangement, with the combustion section (14) comprising: a combustor liner (40) at least partially defining a combustion chamber (50); and a fuel injector assembly (48, 100, 200) comprising a mixing tube (38, 110, 210) having a mixing tube body (112, 212) defining a mixing channel (114, 214). The mixing tube body (112, 212) can have a set of fuel passages (132, 232) terminating in fuel orifices (136, 236) fluidly coupled to the mixing channel (114, 214). A set air flow passages (142, 242) terminating in air outlets (144, 244) can be fluidly coupled to the mixing channel (114, 214). At least some of the air outlets (144, 244) circumscribe a corresponding fuel orifice (136, 236).

A gas turbine engine (10), comprising: a compressor section (12), a combustion section (14), and a turbine section (16) in a serial flow arrangement, with the combustion section comprising: a combustor liner (40) that at least partially defines a combustion chamber (50); and a fuel nozzle assembly (48), comprising: a liquid fuel supply (34) to supply a liquid fuel; a hydrogen fuel supply (36) to supply a gaseous hydrogen fuel; and a fuel nozzle body (49) fluidly coupled with the liquid fuel supply (34) and the hydrogen fuel supply (36) to provide the liquid fuel and the gaseous hydrogen fuel to the combustion chamber (50).

A gas turbine engine (10), comprising: a compressor section (12), a combustion section (14), and a turbine section (16) in a serial flow arrangement, with the combustion section (14) comprising: a combustor liner (40) that at least partially defines a combustion chamber (50); and a fuel nozzle assembly (248, 448, 648), comprising: a liquid fuel supply (34) to supply a liquid fuel; a hydrogen fuel supply (36) to supply a gaseous hydrogen fuel; and a fuel nozzle body (249, 449, 649) defining a fuel nozzle centerline, the fuel nozzle body (249, 449, 649) fluidly coupled with the liquid fuel supply (34) and the hydrogen fuel supply (36) to provide the liquid fuel and the gaseous hydrogen fuel to the combustion chamber (50).