RTX Corporation Patente

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A rotating assembly of a gas turbine engine includes a first rotating component 86; and a second rotating component 66 located radially outboard of the first rotating component 86, relative to an engine central longitudinal axis. A seal assembly 84 is configured to seal between the first rotating component 86 and the second rotating component 66. The seal assembly 84 includes a shuttle 92 located on a radial outer surface 94 of the second component 66 and freely axially movable along the radial outer surface 94 and a piston ring seal 88 retained in the shuttle 92 and engaged with the first rotating component 86. Axial motion of the first rotating component 86 relative to the second rotating component 66 urges movement of the shuttle 92 along the radial outer surface 94 of the second rotating component 66, while the position of the piston ring seal 88 remains stationary relative to the first rotating component 86.

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A tooling assembly for use in forming z-channels in a fibrous ceramic preform includes a mandrel (34) having a first plurality of holes (44) extending into a mandrel body, a first subset of the plurality of holes (44) being through-holes (44<T>) extending completely through the mandrel (44), and a second subset of the holes (44) comprising blind pockets (44<P>), and a plurality of channels (46) extending longitudinally along the mandrel (34) in a direction orthogonal to the first plurality of holes (44). The tooling assembly further includes an outer fixture at least partially enclosing the mandrel (34), the outer fixture including at least one piece (36) comprising a second plurality of holes (38) extending completely through the at least one piece (36), the second plurality of holes (38) being aligned with respective corresponding ones of the first plurality of holes (44) such that a needle can be inserted through each of the second plurality of holes (38) in the at least one piece (36) and into the respective ones of the first plurality of holes (44) in the mandrel (34).

A combustor (56) including: a combustor case (100) defining a plurality of case apertures (105A, 105B); a liner (110) within the combustor case (100) defining a combustion zone (120) and liner apertures (200A, 200B, 200C) through which an airflow flows into the combustion zone (120); a fuel injector (170) having a fuel channel (180) extending through a first case aperture (105A) and the liner (110), and the fuel channel (180) has a nozzle (190) at the combustion zone (120) through which fuel is injected; an igniter (210) for igniting the combustible mixture of fuel and airflow and providing a flame at the nozzle (190); and a flame sensor (220) including: a radio frequency transponder (230), comprising a transmitter-receiver pair, located exterior to the combustor case (100); a horn antenna (250) disposed in the fuel nozzle (190), and a tubular waveguide (240) extending from the radio frequency transponder (230) to the horn (250) via one of the plurality of case apertures (105A, 105B), wherein the flame sensor (220) is configured to perform flame and flow field diagnostics.

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Aircraft propulsion systems (200; 300; 400; 500) including a closed loop-supercritical fluid system (216; 316; 426; 526) having a turbine (214; 314; 428; 528), a cooler heat exchanger (224; 324; 432; 532), a compressor (226; 326; 434; 534), and a recovery heat exchanger (228; 328; 422; 522) arranged along a closed-loop flow path (220; 320; 406; 506) of a supercritical fluid. A shaft (212; 312; 436; 536) is operably coupled to the turbine (214; 314; 428; 528) and configured to be rotationally driven by the turbine (214; 314; 428; 528). A fan (204; 304; 410; 510) is configured to generate thrust, the fan (204; 304; 410; 510) operably coupled to the shaft (212; 312; 436; 536) to be rotationally driven by the shaft (212; 312; 436; 536). A burner (206; 306; 418; 518) is configured to combust a fuel and air from the fan (204; 304; 410; 510) to generate a combusted gas and supply said combusted gas to the recovery heat exchanger (228; 328; 422; 522) of the closed loop-supercritical fluid system (216; 316; 426; 526) and out an exhaust nozzle (210; 310; 424; 524).

A propulsion system (20) for an aircraft includes a hydrogen fuel system (52) suppling hydrogen fuel to a combustor (30) through a fuel flow path . A condenser (80) extracts water (116) from a high energy gas flow (55). The condenser (80) includes a plurality of rotating passages (100) that are disposed within a collector (112) and are configured to rotate about a condenser axis (110) to generate a transverse pressure gradient to direct water (116) out of the high energy gas flow (55) toward the collector (112).

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A turbine component includes a body having a pair of spaced walls, with at least one of the walls for facing a fluid flow when mounted in a gas turbine engine (20). There are a plurality of wall cooling passages (124) having a generally boomerang shape such that a peak apex (140) is spaced from the wall and an indent apex (144) is adjacent to the wall, with the plurality of wall cooling passages (124) having interior sides (146) extending from the peak apex (140) toward the wall to define a corner (142). Outer sides (148) extend from the corners (142) with a component away from the wall and to the indent apex (144). A gas turbine (20) engine is also disclosed.

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A counter weight (76) for use with a rotor module (70), comprises a root portion (90); a main body portion (92) that extends away from the root portion (90); a pair of side portions (93) each extending from opposite sides of the main body portion (92); and at least one plate (100) secured to one of the opposite sides of the main body portion (92) in order to adjust a weight of the counter weight (76). A rotor module (70) comprises: a blade cluster (73) of a plurality of blades (72) secured to a rotor disk (74) of the rotor module (70); and the counter weights (76) secured to the rotor disk (74). The counter weights (76) cause a center of gravity of the plurality of blades (72) to be zeroed out with respect to an axis (O) of the rotor disk (74). The counter weights (76) cause a moment of inertia (Jx) about an X axis of the rotor disk (74) to be equal to a moment of inertia (Jy) about a Y axis of the rotor disk (74) when the rotor module (70) is rotated about the axis (O). A method for zeroing out a center of gravity of a plurality of blades (72) with respect to an axis (O) of a rotor disk (74) the plurality of blades (72) are secured to is also provided.

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A rotor module (70) comprises: a blade cluster (73) of a plurality of blades (72) secured to a rotor disk (74) of the rotor module (70); and counter weights (76) secured to the rotor disk (74). The counter weights (76) cause a center of gravity of the plurality of blades (72) to be zeroed out with respect to an axis (O) of the rotor disk (74). The counter weights (76) cause a moment of inertia (Jx) about an X axis of the rotor disk (74) to be equal to a moment of inertia (Jy) about a Y axis of the rotor disk (74) when the rotor module (70) is rotated about the axis (O). A method for zeroing out a center of gravity of a plurality of blades (72) with respect to an axis (O) of a rotor disk (74) the plurality of blades (72) are secured to is also provided.