General Electric Patente

A gas turbine engine (100) includes a turbomachine (202) comprising a low pressure (LP) spool (204) and a high pressure (HP) spool (206) that rotate about a central axis (102), an electric motor (240) mechanically coupled to the LP spool (204) for selectively rotating the LP spool (204), a starter assembly (210) mechanically coupled to the HP spool (206) for selectively rotating the HP spool (206), and a system controller (270) in operative communication with the electric motor (240) and the starter assembly (210), the system controller (270) being configured for performing an engine startup of the gas turbine engine by: increasing a motor torque of the electric motor (240) to apply torque to the LP spool (204); monitoring one or more engine operating parameters to determine whether the one or more engine operating parameters is within a limit; and increasing the motor torque of the electric motor (240) when the one or more engine operating parameters is within the limit.

The present disclosure generally relates to separating solid particles from an airflow in a gas turbine engine. A system for separating debris includes a first separation device in fluid communication with an inlet flow path of a compressor and a second separation device in fluid communication with an outlet flow path of the compressor and an inlet flow path of a combustor. The first separation device is adapted to remove coarse particles from the airflow. The second separation device is adapted to remove fine particles from the airflow. The course particles have a larger mean particle diameter than the fine particles.

A method for additively manufacturing three-dimensional objects includes generating a laser beam with a laser beam source and splitting the laser beam to form a plurality of beamlets. The plurality of beamlets are collimated via an optical device. The method also includes independently controlling respective beamlets of the plurality of beamlets via respective channels of a multi-channel optical modulator disposed downstream of the optical device to at least one of steer or modulate the respective beamlets. A scanning device disposed downstream of the multi-channel optical modulator scans the respective beamlets over at least a portion of a target plane.

A tool (100) for cleaning a portion of a turbine engine (10) is provided, the tool comprising a rigidizable guide tube (106) including a proximal end (120) and a distal end (122) and configured to be inserted into a portion of the turbine engine (10), a flexible inner tube (104) disposed within a portion of the rigidizable guide tube (106), the flexible inner tube (104) including a proximal end (130) and a distal end (132), a fluid delivery supply (310) coupled to the flexible inner tube (104) for delivering fluid to spray on a surface of an airfoil, and filaments (108) coupled to the distal end (132) (122) of at least one of the flexible inner tube (104) or the rigidizable guide tube (106), wherein when fluid is injected into the flexible inner tube (104), the filaments (108) are configured to agitate the surface of the airfoil.

A method (200) for washing a gas turbine engine (10) includes freezing a cleaning liquid to form solid cleaning granules, compacting the solid cleaning granules to form a compacted cleaning solid (106), and inserting the compacted cleaning solid (106) into the gas turbine engine (10), wherein when the gas turbine engine (10) is started, the compacted cleaning solid (106) liquifies to form a cleaning foam, which spreads through a gas flow path (37) of the gas turbine engine (10) to clean one or more components (22, 24) of the gas turbine engine (10) along the gas flow path (37).

A combustion section (100, 200) for a turbine engine (10) includes a combustor (102, 202) having a combustion chamber (108, 208) and a fuel nozzle (104, 204, 304, 404, 504) for delivering an air-fuel mixture into the combustion chamber (108, 208). The fuel nozzle comprises a body (112, 212, 312, 412, 512) defining a centerline (114, 214, 314, 414, 514) and a central channel (116, 216, 316, 416, 516) extending between an inlet (118, 218) and an outlet (120, 220) that opens to the combustion chamber (108, 208). The central channel (116, 216, 316, 416, 516) includes a compressed air flow passage (122, 222) and a mixer (124, 224) for combining fuel and air. A vane (142, 242, 342, 442, 542) extends from the body (112, 212, 312, 412, 512) and includes an outer wall (144, 244) extending radially between a root (146, 246, 346, 446, 546) and a tip (148, 248, 348, 448, 548) in a spanwise direction (Sd), and axially between a leading edge (150, 250) and a trailing edge (152, 252) in a chordwise direction (Cd). The outer wall (144, 244) includes a plurality of lobes (156, 256, 356, 456, 556) defining a waveform (174, 274, 374, 474, 574) that increases in amplitude (178, 278) in the chordwise direction (Cd) to promote enhanced mixing and uniform combustion.

A recoater assembly (118) is movable across a working surface (116) in a forward stroke and in a return stroke and includes a powder plow (152). One or more actuators (160) are coupled to the powder plow (152) and are actuable to provide vertical movement of the powder plow (152) with respect to the working surface (116). A controller (124) is operable to control the one or more actuators (160) to maintain the powder plow (152) in a retracted position during the forward stroke, lower the powder plow (152) to an extended position when the powder plow (152) is at or near a return area, enable the powder plow (152) to be lifted from the extended position to a partially retracted position a predefined vertical distance from the working surface (116), and maintain the powder plow (152) in the partially retracted position while traversing a portion of the working surface

Gas turbine engines (106) including thermal management systems (200, 300, 400, 500, 600, 700, 764, 800, 808, 900, 1000) with recirculating fuel circuits and related methods are disclosed herein. An example gas turbine engine disclosed herein includes a fuel flow line (202, 708), a fuel inlet (209, 712), a nozzle (212, 714) fluidly coupled to the fuel inlet (209, 712) by the fuel flow line (202, 708), and a recirculation flow circuit (204, 710) coupled to the fuel flow line (202, 708) between the fuel inlet (209, 712) and the nozzle (212, 714), the recirculation flow circuit (204, 710) including a cooler and a heat-producing component (216, 731A) coupled to the fuel flow line (202, 708) downstream of the fuel inlet (209, 712) along the fuel flow line (202, 708).

Methods for manufacturing composite components having complex geometries are provided. In one exemplary aspect, a method includes laying up each of a plurality of laminates to an initial shape with a substantially planar geometry or a gently curved geometry. Then, a laid up laminate is formed to a final shape for each predefined section defined by the composite component to be manufactured. Thereafter, the laminates formed to their respective final shapes are stacked to build up the complex geometry of the composite component. Next, the composite component can be cured and finish machined as necessary to form the completed composite component.

A robotic assembly for servicing equipment, the robotic assembly including an area configured to receive components associated with a workscope of the equipment; an environmental capture device configured to capture information associated with an environment in which the robotic assembly is disposed; and one or more computing devices configured to: locate the equipment in the environment; autonomously navigate the robotic assembly through the environment to the equipment; and autonomously adjust a position of the robotic assembly in response to the workscope.

An electronic component cooling assembly includes an electronic component and a housing defining a slot such that at least a portion of the electronic component is positioned within the slot. The housing further defining an impingement cooling passage fluidly coupled to the slot. The electronic component cooling assembly includes a thermal insert positioned between the electronic component and the housing in the vertical direction, with the thermal insert defining a chamber such that the impingement cooling passage directs a fluid jet into the chamber and onto the electronic component. The thermal insert further defines a cavity spaced apart from the chamber in the longitudinal and/or the transverse direction. Additionally, the electronic component cooling assembly includes a phase change material positioned within the cavity and configured to absorb heat from the electronic component.

In some embodiments, apparatuses and methods are provided herein useful for performing maintenance operations within a gas turbine engine. In some embodiments, a tool (100) performing maintenance operations within a gas turbine engine includes a link assembly (104); an elongated enclosure (102) having a cavity (108) and an enclosure line guide (114); a line assembly (106) inserted through the enclosure line guide (114) and is selectively coupled to the tip link (130) of the link assembly (104); and an actuator apparatus (500) coupled to the link (138) assembly (104) and configured to move the link assembly (104) through the cavity (108) and out of the distal end (148) of the elongated enclosure (102) from a retracted state to an extended state.

Method of manufacturing a woven fabric (400, 408, 500, 504, 520, 600, 604, 700, 702) for a composite component (300, 301) for a gas turbine engine (100). The method includes weaving a plurality of reinforcing fiber tows (202) to form the woven fabric (400, 408, 500, 504, 520, 600, 604, 700, 702). The woven fabric (400, 408, 500, 504, 520, 600, 604, 700, 702) is a three-dimensional woven fabric (400, 408, 500, 504, 520, 600, 604, 700, 702). At least one of a plurality of first fiber tows (210) is arrayed in the thickness direction or a plurality of second fiber tows (220) is arrayed in the thickness direction to form a plurality of fiber layers (222). The method also includes interweaving a plurality of interlocking fiber tows (230) through the plurality of fiber layers (222) to connect the plurality of fiber layers (222) to each other in a thickness direction to form an interlocked region (410), and forming a bifurcated region (420) where adjacent layers of the plurality of fiber layers (222) are free from interconnection by interlocking fiber tows (230) forming a plurality of bifurcated layers (430).

An engine cleaning system (10) for a gas turbine engine (100) includes a first foam cartridge (20) containing a first detergent (22), a second foam cartridge (20) containing a second detergent (22), a first foam generating assembly (30) operably coupled to the first foam cartridge (20), the first foam generating assembly (30) being configured to generate and supply a first cleaning foam (34) to a first cleaning port (146) of the gas turbine engine (100), and a second foam generating assembly (30) operably coupled to the second foam cartridge (20), the second foam generating assembly (30) being configured to generate and supply a second cleaning foam (34) to a second cleaning port (146) of the gas turbine engine (100), wherein the second cleaning foam (34) is different than the first cleaning foam (34).

An assembly for measuring impedance of a battery (24) includes a first current loop (102, 202) including a first capacitor (110, 206) and a second capacitor (112, 208) in series with the first capacitor, a second current loop (104, 204) in parallel with the first current loop, the second current loop including a third capacitor (120, 210), an impedance sensor (106) in electrical communication with the first and second current loops, and a controller module (108) configured to determine an impedance of the battery based on data from the impedance sensor indicating a first current frequency (f1) from the first current loop and a second current frequency (f2) from the second current loop.

A method for performance analysis of an aircraft engine component includes accessing data of an aircraft engine component, the data including a performance signal; detecting a signature of the performance signal; performing a first analysis by an analytical model, based on the data and the performance signature, to determine performance metrics of the aircraft engine component; converting a first performance metric to a second performance metric based on a derivative, the derivative being derived from at least two of the determined performance metrics; determining an analytic for the aircraft engine component based on at least one of the derived derivative or the determined performance metrics; selectively providing a washing recommendation for the aircraft engine component based on the determined analytic; and selectively causing the aircraft engine component to be washed based on the washing recommendation.

A cleaning system (20) for a gas turbine engine (100) includes a detergent supply (22) containing a detergent, the detergent comprising a corrosion-resistance additive (e.g., Gadolinium doped ceria particles, a foam generating assembly (26) operably coupled to the detergent supply (22), the foam generating assembly (26) being configured to generate a cleaning foam using the detergent, and a foam supply line fluidly coupling the foam generating assembly (26) to a cleaning port of the gas turbine engine (100) to supply the cleaning foam into the gas turbine engine (100).

A wash system (300) for washing a gas turbine engine (10) having a compressor section (22, 24), a combustion section (26), and a turbine section (28, 30) includes a detergent reservoir (365) for storing a wash fluid (372), a foam wash module (315) having a foam generating device in fluid communication with the detergent reservoir (365) and configured for generating a flow of foamed wash fluid (376), one or more additive reservoirs (370) for storing at least one additive material (380), and at least one supply line (378) in communication with the foam wash module (315) and the one or more additive reservoirs (370). The at least one additive material (380) includes a self-propelling material, an exothermic material, or both the self-propelling material and the exothermic material. The at least one supply line (378) supplies the foamed wash fluid (376) and the at least one additive material (380) to the gas turbine engine (10).

A method for cleaning a component (202, 302, 702) of a gas turbine engine (200), the component (202, 302, 702) disposed within an engine core, the gas turbine engine (200) defining one or more access ports within the engine core, the method comprising subjecting the component (202, 302, 702) to two or more transient thermal cycles, wherein a surface temperature of the component (202, 302, 702) changes by at least 293 degrees kelvin (K) within a transient thermal cycle, and cleaning the component (202, 302, 702) to remove at least a portion of a contaminant deposit (204) from the component (202, 302, 702).

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-air mixer (48) fluidly coupled with the combustion chamber (50) and comprising: a mixing tube (100) defining a mixing channel (102) including a mixing tube inlet (104), a mixing tube outlet (106) fluidly coupled with the combustion chamber (50), and a mixing channel centerline, the mixing tube inlet (104) including a fuel inlet (110) to receive fuel and an air inlet (112) to receive air; a first set of air jets (130) fluidly coupled with the mixing channel (102) and axially spaced downstream of the mixing tube inlet (104); and a second set of air jets (132) fluidly coupled with the mixing channel (102).

A system for cleaning gas turbine engines (10) includes at least one particle cleaning medium reservoir (106) for storing at least one particle cleaning medium. The at least one particle cleaning medium includes at least one of a first plurality of organic particles or a second plurality of organic particles, with each of the first plurality of organic particles being dissolvable, and with each of the second plurality of organic particles having fibers (F1). The system further includes a delivery system (108) being configured to direct the at least one particle cleaning medium from the at least one particle cleaning medium reservoir (106) towards one or more components of a gas turbine engine (10) for cleaning the one or more components of the gas turbine engine (10).

A system (200) for reducing wear during a wash of a gas turbine engine (10), the system comprising:an accessory gearbox (45);a transfer gearbox (46) drivingly and fluidly coupled to the accessory gearbox (45); anda lubrication system (100) fluidly coupled to the accessory gearbox (45) and the transfer gearbox (46), the lubrication system (100) comprising:a lubricant reservoir (102);a supply circuit (106) to supply lubricant (104) from the lubricant reservoir (102) to the accessory gearbox (45) and the transfer gearbox (46);a return circuit (108) to return lubricant (104) from the accessory gearbox (45) and the transfer gearbox (46) to the lubricant reservoir (102); anda wash system connector (101) coupled to a wash system (80) and fluidly coupled to the supply circuit (106) upstream of the accessory gearbox (45) and the transfer gearbox (46), the wash system connector (101) configured to supply a mixture of air and lubricant from the wash system (80) to the accessory gearbox (45) and the transfer gearbox (46).

A gas turbine engine includes a turbomachine having a compressor section, a combustion section, and a turbine section arranged in serial flow order. The turbomachine defines an engine inlet to an inlet duct, a fan duct inlet to a fan duct, and a core inlet to a core duct. The primary fan is driven by the turbomachine, and a secondary fan is located downstream of the primary fan within the inlet duct. One or more actuation devices operably associated with the fan duct, the one or more actuation devices actuable to increase or decrease an exit area of the fan duct.

A system for reducing wear during a wash of a gas turbine engine (10) is provided. The system includes an accessory gearbox (45), a transfer gearbox (46) drivingly coupled to the accessory gearbox (45), and a lubrication system (100) fluidly coupled to the accessory gearbox (45) and the transfer gearbox (46). The lubrication system (100) includes a supply circuit (106) for supplying lubricant (104) from a lubricant reservoir (102) to the accessory gearbox (45) and the transfer gearbox (46). The lubrication system further comprises a return circuit (108) for returning the lubricant (104) from the accessory gearbox (45) and the transfer gearbox (46) to the lubricant reservoir (102). The lubrication system further comprises a wash bypass line (168) extending from the return circuit (108) to the supply circuit (106).

A wash system for a gas turbine engine, with an inlet assembly (102, 202, 302, 402) comprising a housing defining a mixing chamber (134, 234, 334, 434) with a porous structure (136, 236, 336, 436) therein, a first fluid conduit (104, 204, 304, 404) in fluid communication with a gas source (118), a second fluid conduit (106, 206, 306, 406) in fluid communication with a detergent source (120) and a third fluid conduit (121, 221, 321, 421) in fluid communication with a liquid source (119); a delivery nozzle (108, 208, 308, 408) coupled to the inlet assembly (102, 202, 302, 402); a plurality of valves (116A, 116B, 116C, 216A, 216B, 216C, 316A, 316B, 316C, 416A, 416B, 416C) to control fluid flow from the first fluid conduit (104, 204, 304, 404), the second fluid conduit (106, 206, 306, 406), and the third fluid conduit (121, 221, 321, 421) into the delivery nozzle (108, 208, 308, 408); and a controller configured to actuate the plurality of valves (116A, 116B, 116C, 216A, 216B, 216C, 316A, 316B, 316C, 416A, 416B, 416C) to cause the wash system to switch between a foam washing mode and a liquid washing mode, wherein the inlet assembly delivers foam to the delivery nozzle (108, 208, 308, 408) in the foam washing mode and delivers liquid from the liquid source (119) to the delivery nozzle (108, 208, 308, 408) in the liquid washing mode.

Provided herein are systems and methods for supplying a gas to one or more portions of an engine (10) during an engine cleaning operation to limit the ingress of cleaning fluid. The systems and methods can also be used to supply the gas to the engine (10) after a cleaning operation to remove residual cleaning fluid. In some embodiments, the methods include pressurizing a bleed system (60) to limit ingress of cleaning fluid or dry the bleed system (60). The methods may use a portion of a reverse bleed apparatus (90) associated with the bleed system (60) to supply the gas to the bleed system (60). In addition, an external gas supply system (130, 150, 166, 220, 240) may be coupled to a bleed duct (74) to supply the gas. In some embodiments, the methods include pressurizing a sump system (168) of the gas turbine engine (10) to limit ingress of cleaning fluid into the sump system (168) during the cleaning operation.

A fuel injector (100) is provided for a turbine engine (10) comprising a compression section (12), a combustion section (14), and a turbine section (16) is serial flow arrangement. The fuel injector (100) includes an outer wall (102) surrounding an injector passage (104) and defining a longitudinal injector axis (106). The injector passage (104) exhausts at an injector outlet (110) and defines a flow direction (FD) through the injector passage (104). A first converging portion (130) is provided in the outer wall (102). A swirler assembly (124) configured to provide a swirled airflow to the injector passage (104) upstream of the first converging portion (130) relative to the flow direction (FD). A first set of fuel passages (140) are provided in annular arrangement about the outer wall (102) fluidly exhausting to the injector passage (104) downstream of the first converging portion (130) relative to the flow direction (FD).

A wash system (300A, 300B, 300C) for washing a gas turbine engine (10) having a compressor section (22, 24), a combustion section (26), and a turbine section (28, 30) includes a detergent reservoir (365) for storing a wash fluid (372), a foam wash module (315) in fluid communication with the detergent reservoir (365) for generating a flow of foamed wash fluid (376), a particulate reservoir (370) for storing solid particulates (380), and at least one supply line (378) fluidly coupled to the foam wash module (315) and the particulate reservoir (370) for supplying a mixture of the foamed wash fluid (376) and the solid particulates (380) to the gas turbine engine (10).

A gas turbine engine (100) is provided. The gas turbine engine includes a turbomachine (120, 506) including a compressor section (512), a combustion section (514), and a turbine section (516) arranged in serial flow order, the turbomachine (120, 506) defining an engine inlet (182) to an inlet duct (180), a fan duct inlet (176) to a fan duct (172), and a core inlet (124) to a core duct (142); a primary fan (152) driven by the turbomachine (120, 506); and a secondary fan (184) located downstream of the primary fan (152) within the inlet duct (180), the gas turbine engine (100) defining a thrust to power airflow ratio between 3.5 and 100 and a core bypass ratio between 0.1 and 10, wherein the thrust to power airflow ratio is a ratio of an airflow through a bypass passage (194, 522) over the turbomachine (120, 506) plus an airflow through the fan duct (172) to an airflow through the core duct (142), and wherein the core bypass ratio is a ratio of the airflow through the fan duct (172) to the airflow through the core duct (142).

A wash system (300A, 300B) for washing a gas turbine engine (10) includes a particulate reservoir (405) for storing a cleaning material and at least one particulate supply line (410) coupled to the particulate reservoir (405) and configured to supply the cleaning material to an engine inlet of the gas turbine engine (10). The cleaning material is at least partially in a solid state and comprises a detergent. Each particulate (450) of the cleaning material comprises a first plurality of particles (455) and a second plurality of particles (460) different from the first plurality of particles (455). The first plurality of particles (455) at least partially surrounding the second plurality of particles (460).

A system (200) includes an accessory gearbox (45), a transfer gearbox (46) that is drivingly and fluidly coupled to the accessory gearbox (45), and a lubrication system (100) that is fluidly coupled to the accessory gearbox (45) and the transfer gearbox (46). The lubrication system (100) includes a lubricant reservoir (102) that is fluidly connected to the transfer gearbox (46) and the accessory gearbox (45). A solenoid valve (140) is in fluid communication with the lubricant reservoir (102). The solenoid valve (140) is selectively actuatable upon providing an electrical current (101) from a power source (88, 91) between an open position that allows lubricant (104) to flow from the lubricant reservoir (102) to the accessory gearbox (45) and the transfer gearbox (46) and a closed position that prevents the lubricant (104) to flow from the lubricant reservoir (102) to the accessory gearbox (45) and the transfer gearbox (46).

A system for cleaning gas turbine engines includes a cleaning medium reservoir (104) for storing a cleaning medium. The system further includes a delivery system (102A) configured to direct the cleaning medium from the cleaning medium reservoir (104) towards one or more components of a gas turbine engine (10) for cleaning the one or more components of the gas turbine engine (10). The system additionally includes a removal system (102B) configured to apply a vacuum pressure to the gas turbine engine (10) for moving the cleaning medium through the gas turbine engine (10).

A system 100 for wash time estimation for aircraft components 300, 310 includes a processor 220 and a memory 230. The memory 230 includes instructions which, when executed by the processor 220, cause the system 100 at least to perform capturing an image of a surface of a stud 308 disposed in a portion of an aircraft component 300, 310; determining a volume of particulate matter 314 on a surface of the stud 308 based on the captured image; providing, by an analytic model, a wash parameter based on the volume of particulate matter 314; and recommending a wash based on the determined wash parameter.

A cleaning system (20) for a gas turbine engine (100) includes a foam supply unit (26) for generating cleaning foam (194) that is injected into the gas turbine engine (100), and an air blower (180) that is operably coupled to a core flowpath (121) of the gas turbine engine (100), the air blower (180) being configured to draw a flow of air (182) through the core flowpath (121) at a flow rate that induces rotation of the gas turbine engine (100) at a target rotational speed.

A system for querying a federated data store includes a metadata knowledge graph describing the contents and relationships among one or more underlying data stores, an interactive user interface receiving requests from a data consumer, a predefined constrainable query ('nodegroup') store containing predefined constrainable queries that define data subsets of interest across one or more of the underlying data repositories, a knowledge-driven querying layer generating and executing queries against the federated data store and merging responsive results, a scalable analytic execution layer receiving the search results from the federated data store and applying machine learning/artificial intelligence techniques to analyze the results, and a user interface presenting visualizations of raw or analyzed results to the consumer. A method and a non-transitory computer-readable medium are also disclosed.

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), and loading area (840, 1090) (LA) are related as follows: EQ 1 = IBR m πm * LA m 2 0.00001 m 2 * IBFA m 2 0.000001 m 2 2 <img class="EMIRef" id="07eadbc3-2700-4c32-9450-2d87ecd618c5-ia01" /> , wherein 0.0000002 meters squared (m<2>) ≤ IBFA ≤ 0.0000019 m<2>, 0.213 m ≤ IBR ≤ 0.221 m, 0.0000147 m<2> ≤ LA ≤ 0.0000153 m<2>, and 0.006 ≤ Equation 1 ≤ 0.393.

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). 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).

A gas turbine engine (10) including a compressor section (12), a combustion section (14), and a turbine section (16) in a serial flow arrangement, with the combustion section (14) having a fuel nozzle (48, 104, 204, 304, 404, 504, 604) including a fuel nozzle body (38, 112, 212, 312, 412, 512, 612) defining an axis (114, 214, 314, 414, 514, 614) and having an inner surface (115, 215, 315, 415, 515, 615) defining a channel (116, 216, 316, 416, 516, 616) fluidly coupled to a combustion chamber (50, 108, 508, 608), a support matrix (126, 226, 326, 426, 526, 626) located within the channel (116, 216, 316, 416, 516, 616) comprising a plurality of segments (154, 156, 160, 162, 266, 268, 370, 372) which intersect each other when viewed from aft, and a set of vortex generators (132, 232, 332, 432, 532, 632) located on the support matrix (126, 226, 326, 426, 526, 626).

Provided herein are compositions and methods for removing a coating, such as an environmental barrier coating (EBC), from a substrate. By some approaches, a method for removing a coating from a substrate comprises contacting the coating with a salt, the salt being a solid. The method also involves heating the coating and the salt to a temperature at or above the melting point of the salt to convert the coating to a dissolvable reaction product by fusion of the coating with the salt. After forming the dissolvable reaction product, the product is then dissolved using an acid to remove at least a portion of the coating from the substrate.

An aircraft comprising a fuselage and an unducted turbine engine. The fuselage having a divot with an upstream edge and a downstream edge. The divot is defined by a straight reference line having a length (L) and a maximum depth (h) relative to the straight reference line. The unducted turbine engine having an engine core, a nacelle, and a set of blades. A first flow ratio (FR1) is equal to: h L <img class="EMIRef" id="07e75abe-22b9-4a70-b926-77b750086a02-ia01" /> .