RTX Corporation Patente

A method of operating an additive manufacturing system (20) utilizing a layer-by-layer thermo-mechanical analysis system (68) includes the steps of pre-modeling a structure or workpiece (38) into a plurality of slices (42) stacked from a bottom slice (42C) to a top slice (42A) of the plurality of slices (42), assuming a second slice (42B) modeled immediately below a first slice (42A) is a rigid substrate, and run a response force analysis for the first slice (42A). The analysis system may reiterate this process for each slice (42A-42C) calculating a force distribution map (72) for each slice (42A-42C), and adding the force distribution maps (72) to a total response force map (70) of the entire structure (38). The additive manufacturing system (20) may then use the total response force map (70) to optimize material properties for selected regions of each slice (42A-42C) to a sacrificial support structure (74S) and/or the final product.

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A component for a gas turbine engine (20) includes an airfoil shell (92; 192) having an internal cavity (84) and extending between a radially inner platform (70) and a radially outer platform (72). The component also includes a spar (86; 186) disposed within the internal cavity (84). The airfoil shell (92; 192) has a greater resistance to heat than the spar (86; 186), and the spar (86; 186) has a greater strength than the airfoil shell (92; 192). A radiant heat shield (90; 190; 290; 390) circumscribes the spar (86; 186) within the internal cavity (84). The radiant heat shield (90; 190; 290; 390) is enclosed within the airfoil shell (92; 192) and has a reflectance to radiant heat that is greater than or equal to that of the spar (86; 186).

A heat shield vane support (300) may have a structure including at least one of a cylindrical (330) and a frustoconical (332) portion with an outboard diameter (308) and an inboard diameter (306) extending about an axis (301) and extending axially between a forward face (302) and an aft face (304) and may comprise a flange (310) extending circumferentially about the aft face (304) radially inward from the inboard diameter (306), a plurality of anti-rotation features (314) distributed circumferentially about the inboard diameter (306) extending radially inward from the inboard diameter (306) and extending axially with respect to the cylindrical structure between a forward edge (313) and an aft edge (315) defining a length L, and a plurality of teeth (312) distributed circumferentially about the forward face (302) and extending axially from the forward face (302).

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A guide vane retention system (100) for a gas turbine engine (20). The system includes an outer diameter shroud (86) defining an aperture. Also included is a guide vane (62) having a radially outer end (87) extending through the aperture, the guide vane (62) defining a slot (104) proximate the radially outer end (87) and positioned radially outwardly of the outer diameter shroud (86) in an installed condition of the guide vane (62). Further included is a clip (102) disposed within the slot (104) of the guide vane (62). The clip (102) includes a looped end (110). The clip (102) also includes a first leg (112) extending away from the looped end (110) to a first free end (116). The clip (102) further includes a second leg (114) extending away from the looped end (110) to a second free end (118).

A bowed rotor prevention system (42) for a gas turbine engine (10) includes a core turning motor (44) operable to drive rotation of an engine core of the gas turbine engine (10). The bowed rotor prevention system also includes a full authority digital engine control (FADEC) (102) that controls operation of the gas turbine engine (10) in a full-power mode and controls operation of the core turning motor (44) to drive rotation of the engine core using a reduced power draw when the FADEC (102) is partially depowered in a low-power bowed rotor prevention mode.

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

A bearing assembly for a gas turbine engine includes a bearing housing (204) having a first bearing housing surface (205) and a second bearing housing surface (206). The bearing housing surfaces (205, 206) are substantially axially parallel. A the first bearing housing surface (205) has a first and second piston ring groove (212, 213) with a corresponding piston ring (214, 215) in each groove. A damper (210) includes a first outer surface (209) radially adjacent to and opposing the first bearing housing surface (205) and a second outer surface (216) radially adjacent to and opposing the second bearing housing surface (206). The second outer surface (216) includes a third and fourth piston ring groove (217, 219) with a corresponding piston ring (220, 222) disposed in each groove. The bearing housing (204) includes a damping fluid passage (224) configured to provide damping fluid to a gap (211) radially between the first outer surface (209) and the first bearing housing surface (205) and to a gap (223) radially between the second outer surface (216) and the second bearing housing surface (206). A method for maintaining pressure in a squeeze film bearing damper is also provided.

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A gas turbine engine (100) comprises a compressor section having a downstream most end (113) and a cooling air tap (120) at a location spaced upstream from the downstream most end (113). The cooling air tap (120) is passed through at least one boost compressor (118) and at least one heat exchanger (124, 130), and then passed to a turbine section to cool the turbine section, the boost compressor (118) being controlled to provide a desired pressure to the turbine section.

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

A method is provided that involves a component (20) including a first fastener aperture (36) that extends through the component (20). During the method, the component (20) is machined to enlarge the first fastener aperture (36) to provide an enlarged aperture (54). The component (20) is friction plug welded to plug the enlarged aperture (54) with friction plug welded material (56). A second fastener aperture (36B) is machined in the friction plug welded material (56), where the second fastener aperture (36B) extends through the component (20).

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

A lower pressure tap (46) is connected to a first heat exchanger (48) to be cooled by cooling air, and then to a selection valve (52). The selection valve (52) selectively delivers the lower pressure tap air to a boost compressor (44). The lower pressure tap air downstream of the boost compressor (44) is connected to cool the at least one turbine (32). The selection valve (52) also selectively delivers a portion of the lower pressure tap air across a first cooling turbine (78), and to a line (79) associated with an air delivery system for a cabin (80) on an associated aircraft. A portion of the air downstream of the first cooling turbine (78) is connected to a second cooling turbine (84), and air downstream of the second cooling turbine (84) is connected for use in a cold loop (88).

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

A gas turbine engine (10) includes a first vane (328a-1) of a turbine section (221), a second vane (328b-1) of the turbine section (221), an impingement plate (352) that includes a first impingement hole (352a) and a second impingement hole (352a), and a pipe that supplies a fluid. The first vane (328a-1) includes a first platform (232-1). The second vane (328b-1) includes a second platform (232-1) and a baffle (356b-1). A first portion of the fluid (352b) traverses the first impingement hole (352a) and impinges on a first surface (408a) of the first platform (232-1). A second portion of the fluid (352b) traverses the second impingement hole (352a) and impinges on a second surface (408b) of the second platform (232-1). A third portion of the fluid (356a) bypasses the impingement plate (352) and enters the baffle (356b-1).

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Zusammenfassung wird geladen …

A tangential on-board injector (TOBI) for a gas turbine engine is provided. The TOBI includes a body (212) having an entrance and an exit, the body having an outer curved wall and defining an air passageway (214) between the entrance and the exit, and a purge cavity (230) configured on an exterior surface of the body with a purge cavity entry (228) fluidly connecting the air passageway (214) to the purge cavity (230). The purge cavity (230) is configured such that particles in an airflow through the passageway (214) will enter the purge cavity (230).

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Combustors of gas turbine engines having a combustor shell (330; 430; 530; 630; 730; 830; 930; 1030) having a first end (342; 942) and a second end (344; 1044) opposite the first end, a first securing element (346; 946) positioned at the first end of the combustor shell, a second securing element (348; 1048) positioned at the second end of the combustor shell, a plurality of high temperature material panels (340; 440; 540; 640; 740; 840; 940; 1040) fixedly secured by the first securing element (346; 946) at the first end (342; 942) and the second securing element (348; 1048) at the second, wherein a panel gap (454; 554; 654) is formed between edges of adjacent high temperature material panels of the plurality of high temperature material panels, and a seal divider (450; 550; 650) extending from the first end (342; 942) to the second end (344; 1044) and positioned on the combustor shell and arranged to seal the panel gap (454; 554; 654) between adjacent first and second high temperature material panels.

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

According to an aspect, a system includes a starter air valve (116) in fluid communication with an air turbine starter (120) to drive motoring of a gas turbine engine responsive to a compressed air flow (108) from a compressed air source (114). The system also includes a variable-position electromechanical device (104) operable to adjust positioning of the starter air valve (116) and a discrete-position electromechanical device (106) operable to adjust positioning of the starter air valve (116) and limit a motoring speed of the gas turbine engine below a resonance speed of the gas turbine engine responsive to a pulse width modulation control based on a failure of the variable-position electromechanical device (104).

Zusammenfassung wird geladen …