RTX Corporation Patente

A gas turbine engine case structure includes the unitary combination of: a case wall having an inner surface and an outer surface; at least one fuel inlet (264); circumferentially-distributed fuel injectors protruding inward from the case wall and having an outlet; and a fuel plenum fluidically between the at least one fuel inlet (264) and the fuel injectors and configured so that each inlet (264) of the at least one inlet (264) is coupled to feed multiple of the fuel injectors. The fuel plenum alternatingly passes forward and aft.

An assembly is provided for an aircraft propulsion system. This assembly includes a carrier (90), a first gear system (86), a second gear system (88) and a lock device (106). The first gear system (88) includes a sun gear (92), a first ring gear (94) and a plurality of first intermediate gears (96) between and meshed with the sun gear (92) and the first ring gear (94). Each of the first intermediate gears (96) is rotatably mounted to the carrier (90). The second gear system (88) includes a second ring gear (102) and a plurality of second intermediate gears (104) meshed with the second ring gear (102). Each of the second intermediate gears (104) is rotatably mounted to the carrier (90). The lock device (106) is configured to lock rotation of the carrier (90) about an axis (28) during a first mode. The lock device (106) is configured to unlock rotation of the carrier (90) about the axis (28) during a second mode.

A fluid injection system (200; 300; 400) for a gas turbine engine (110) may comprise a fluid injector (160) configured to inject a fluid into an exhaust flow exiting a turbine section (28) of the gas turbine engine (110). The fluid injector (160) may be coupled to a turbine exit guide vane (150) located at a forward end of an exhaust system of the gas turbine engine (110). The fluid may decrease a temperature of the exhaust flow exiting the turbine section (28) and/or increase a thrust of the gas turbine engine (110).

An electronics pressurization system (100) includes a compressed air source (125) providing pressurized air, a manifold (115), and a case (105a; 105b) containing a first high-voltage electrical system, the case (105a; 105b) comprising a pressurizable vessel and a vent hole (106a, 106b, 106c, 106d; 106e). The compressed air source (125) is connected to the manifold (115), and the manifold (115) is connected to the case (105a; 105b). The compressed air source (125) provides a flow of pressurized air, via the manifold (115) to the case (105a; 105b) to maintain an air pressure in the case (105a; 105b) at a higher-than-ambient pressure.

A method for repairing a gas turbine engine component (1000), comprising: providing at least one gas turbine engine component (1000) comprising a surface area having a portion comprising a first size and corrosion; removing the corrosion from the portion to form a portion free of corrosion and comprising a second size; treating the portion free of corrosion and comprising a second size to form a portion free of corrosion; disposing an amount of at least one anti-corrosion agent on the treated portion free of corrosion to form an anti-corrosion agent-coated treated portion free of corrosion; and removing at least a portion of the anti-corrosion agent from the anti-corrosion agent-coated treated portion free of corrosion to form a portion free of corrosion comprising the first size and a residual amount of the anti-corrosion agent.

A ceramic matric composite (CMC) component (100) for a gas turbine engine, such as blade outer air seal (BOAS) (101), may be shielded from thermal stress. The CMC component (100) includes a first surface (110) configured for exposure to a hot gas stream (112), a second surface (120) configured for exposure to a cold gas stream (122), an edge surface (130) of the CMC component (100) disposed between and connecting the first (110) and second (120) surfaces. At least one coating layer (140) is disposed on the first surface (110) and wrapped over the edge surface (130) and the second surface (120), wherein the at least one coating layer (140) disposed on the second surface (120) has a lower thermal conductivity than the CMC component (100) so as to reduce a thermal gradient between the cooled first surface (110) and the heated second surface (120).

A method can comprise dividing a heat exchanger design into a plurality of modules (10), the plurality of modules (10) arranged in a grid, each module in the plurality of modules (10) including: a first fluid conduit (161) defining an inlet (1611), an outlet (1612), and a heat-transfer surface (602), and a first flow direction, and a second fluid conduit (162) defining a second inlet, a second outlet, a second heat-transfer surface, and a second flow direction, the second flow direction different from the first flow direction; and determining a heat-transfer augmenter arrangement (1161) for the first fluid conduit (161) and the second fluid conduit (162) of each module in the plurality of modules (10) based on a stress threshold of the module (16) in the plurality of modules (10).

An artificial diffraction inspection sample including a sample component having a component body with an exterior surface; and a complimentary flat bottom hole formed in the component body, the complimentary flat bottom hole comprising a complimentary bottom surface with a complimentary incidence angle, the complimentary bottom surface being rotated around an acoustic path vector; wherein the complimentary bottom of the complimentary flat bottom hole represents an axial-radial crack.

A system (20) for inspecting a gas turbine engine component (23) includes an inspection environment (22) operable to access feature information associated with a component design (36). The feature information includes a unique identifier (40) assigned to a respective geometric feature (38) of the component design (36). The environment (22) is operable to access one or more inspection criterion (47) associated with the unique identifier (40). The environment (22) is operable to access process information (45) obtained during manufacturing of a physical instance of the respective geometric feature (38) according to one or more manufacturing parameters (48). The environment (22) is operable to generate, using a machine learning model (50), a prediction of whether the physical instance of the respective geometric feature (38) meets the one or more inspection criterion based (47) on the process information (45) . A method for inspecting a component is also disclosed.

A turbine engine rotor has a central shaft and a disk stack having a plurality of disks encircling the shaft. A seal (100) has a first member (102) having, in central axial cross-section: a rearwardly-open channel (240) receiving a portion (260) of one of the disks; a sleeve (232) extending rearward from the channel (240) and between the disk and the shaft; and a portion (230) extending radially inward from the sleeve (232) and having a forward surface (234). The seal (100) has a second member (101) having, in central axial cross-section: a forward-facing surface (116) contacting an aft-facing surface (224) of the shaft (130); and a rear-facing surface (109) contacting the forward surface (234) of said portion (230) extending radially inward from the sleeve (232).

A gas turbine engine is provided and includes a first fan blade (201₁) including a suction surface, a second fan blade (201₂) comprising a pressure surface (215) and neighboring the first fan blade (201₁) and a throat region (220) interposed between the suction surface (216) of the first fan blade (201₁) and the pressure surface (215) of the second fan blade. The throat region (220) includes a passage throat (221) located at a minimum distance between the pressure and suction surfaces. The first and second fan blades are configured such that a pre-compression region is defined in the throat region (220) ahead of the passage throat. Each of the first and second fan blades includes a mean camber line (MCL) defining a flattened suction surface.

A turbine engine rotor blade (76) is provided that has an airfoil (78) that includes a side wall (98) with exterior and interior wall surfaces. The airfoil (78) includes a first wall channel (104) that is defined by a base surface (110) that is a portion of the side wall interior surface and first and second interior wall surfaces (112, 114). A first interface surface (116) extends between the base surface (110) and the first interior wall surface (112). A second interface surface (118) extends between the base surface (110) and the second interior wall surface (114). A third interface surface (120) extends between the first interior wall surface (112) and the second interior wall surface (114). The base surface (110) includes base surface first and second segments (110A, 110B), and a peak interface surface (122) extending between the base surface first and second segments (110A, 110B). The first interface surface (116) is disposed closer to the side wall exterior surface (92) than both the second interface surface (118) and the peak interface surface (122).

A method for assembling a plug assembly (100) for plugging one or more ports of a gas turbine engine including that a connector body (160) and a cover (150) operably connected to the connector body (160) are inserted into a sheath through-passage (141) of a sheath (140). The connector body (160) including a rotatable joint (169) operably connecting the cover (150) to the connector body (160). The method also includes that a biasing mechanism (192) configured to apply a force to the cover (150) is installed, the cover (150) is rotated relative to the connector body (160) via the rotatable joint (169), a top housing (180) is slid over the biasing mechanism (192) such that the biasing mechanism (192) abuts a bottom end (182) of the top housing (180) or is located in a cavity (186) defined within the top housing (180), and the top housing (180) is secured together with the sheath (140).

A heat exchanger fairing including a fairing wall comprising a flexible region configured to change at least one dimension responsive to thermal expansion or contraction of a heat exchanger.

A fuel nozzle thermal management system including a fan fluidly coupled with at least one of a fuel nozzle, a fuel manifold, an external fuel supply, and fuel drain lines proximate a combustor.

A coated substrate is described for use as part of a jet engine component which includes a substrate and a coating system. The substrate can be a ceramic matrix composite or a superalloy substrate. The coating system is applied to the substrate by thermal spraying such as APS. The coating system includes a layer containing as hafnium silicate (hafnon), zirconium silicate (zircon), a rare earth phosphate (REPO₄), rare earth oxides (RE₂O₃), alumina, an aluminosilicate, rare earth-stabilized zirconia, and HfO₂-SiO₂-rare earth (RE) oxide, hafnia (HfO₂) stabilized (partially or fully) by the addition of another component, zirconia (ZrO₂) coating stabilized (partially or fully) by the addition of another component, a rare earth zirconate (RE₂Zr₂O₇), a rare earth hafnate (RE₂Hf₂O₇), and combinations thereof, wherein RE is Sc, Y, La, Ce, Pr, Nd, Pm, Sm, Eu, Gd, Tb, Dy, Ho, Er, Tm, Yb, or Lu. To mitigate crack formation during amorphous-crystalline phase transformation, the layer further contains 0.1 to 10 wt.% of a metal oxide selected from Al₂O₃, SiO₂, Nb₂O₃, MgO, CaO, SrO, BaO, and combinations thereof, as a sintering agent.

A nut strip assembly is provided that includes a carrier and a plurality of nuts. The carrier includes a base, a plurality of first tabs and a plurality of second tabs. The base is configured with a plurality of apertures. Each of the first tabs is connected to and project out from a first side of the base. Each of the second tabs is connected to and projects out from a second side of the base. The nuts are secured to the carrier. Each of the nuts are aligned with a respective one of the apertures. Each of the nuts is captured between the base and a respective one of the first tabs. Each of the nuts captured between the base and a respective one of the second tabs.

A method of forming a heat exchanger includes the steps of forming an inlet header (102) having a hollow interior formed by a wall, forming heat transfer tubes (108) of a diameter smaller than a diameter of the inlet header (102) and extending away from the inner wall of the inlet header (102), the heat transfer tubes (108) having a hollow interior, with the hollow interior of the inlet header (102) being blocked from communication with the hollow interior of the heat transfer tubes (108), forming an outlet header having a hollow interior formed by a wall, the outlet header formed at an opposed end of the inlet header (102), and the hollow interior of the outlet header being in communication with the hollow interior of the heat transfer tubes (108), cutting access opening through the wall of the inlet header (102) at a location opposed to the heat transfer tubes (108), and cutting through an opposed side of the wall of the inlet header (102) to form orifices (140) to communicate the interior of the inlet header (102) to the interior of the heat transfer tubes (108), then closing the access openings. A heat exchanger arrangement is also disclosed.

A composite fan blade (58) includes a plurality of uniplies (74) stacked together to define the blade (58); and at least one braided (78) or woven (76) ply between two plies of the plurality of uniplies (74). Z-pins (96) can be incorporated into the fan blade (58. A method for making a composite fan blade is also disclosed.

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A process for solid lubricant filler powder used in abradable coating manufacture comprising mixing (12) a bentonite clay and a hexagonal boron nitride powder to form a mixture of the bentonite clay and the hexagonal boron nitride powder; consolidating (14) the bentonite clay and the hexagonal boron nitride powder to form a composite material; heat treating (16) the composite material to at least 500 degrees centigrade; breaking up (18) the composite material into a variety of sizes; and segregating (20) the composite material to produce a final product of free flowing, low dust powder of composite hexagonal boron nitride and calcined bentonite.

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A propulsion system for an aircraft including a gas generating core engine (70) that generates an exhaust gas flow (52) that is expanded through a turbine section (32). A power turbine (38) engine is forward of the core engine (70). A propulsor (22) is coupled to the power turbine (38). A hydrogen fuel system (96) supplies hydrogen fuel to the combustor (30) through a fuel flow path (80). A condenser (62) extracts water (54) from the exhaust gas flow (52). An evaporator (64) receives a portion of the water (54) that is extracted by the condenser (62) and generates a steam flow (56). The steam flow (56) is injected into the core flow path (C) upstream of the turbine section (32). An inlet duct (66) communicates an inlet airflow (48) to the compressor section (24) aft at a location aft of the turbine section (32). An exhaust duct (68) routes exhaust gas flow (52) through the condenser (62) and into thermal communication with a water flow (86) of water (54) extracted in the condenser (62) in the evaporator (64).

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A composite laminate comprising a substrate (54) having a hot side (60) and a cold side (58) opposite said hot side (60); a thermal barrier coating (62) coupled to said hot side (60) of said substrate (54); and a reflective coating (76) disposed on said thermal barrier coating (62).

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