RTX Corporation Patente

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An airfoil piece (62) includes at least one inner ceramic matrix composite ply (100) which defines an internal cavity (72) of an airfoil section (70) and first and second collar projections (74). At least one inner ceramic matrix composite ply (100) is continuous through the airfoil section (70) and first and second collar projections (74), first and second platforms (66, 68) at the first and second ends of the airfoil section (70). The first and second collar projections (74) extend radially past the first and second platforms (66, 68), respectively. A vane (60) for a gas turbine engine (20) and a method of assembling a vane (60) are also disclosed.

An airfoil (60) includes an airfoil wall (62) that defines a leading end (62a), a trailing end (62b), a first side (62c), and a second side (62d). Radially-extending ribs (64) partition the interior cavity of the airfoil into first and second cooling channels (72) and a radial cooling passage that is situated between the first and second cooling channels. The cooling channels extend to respective first and second channel ends. A turn channel connects the first and second channel ends. The turn channel splits at the first channel end into first and second channel legs (80, 82) such that there is a region between the first and second channel legs. The channels legs merge at the second channel end. The radial cooling passage extends through the region between the first and second channel legs.

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A turbine engine (20) is provided that includes a first rotating assembly (80), a bearing (90D), a turbine engine core (34) and a flowpath (52). The first rotating assembly (80) includes a propulsor rotor (66), a power turbine rotor (67) and a power turbine shaft (78) coupled with and axially between the propulsor rotor (66) and the power turbine rotor (67). The bearing (90D) rotatably supports the first rotating assembly (80) at an intermediate location (102) along the power turbine shaft (78). The turbine engine core (34) includes a second rotating assembly (84, 88) and a combustor (122). The second rotating assembly (84, 88) includes a core compressor rotor (68, 69) and a core turbine rotor (70, 71). The power turbine rotor (67) is arranged axially between the propulsor rotor (66) and the second rotating assembly (84, 88). The flowpath (52) extends sequentially across the core compressor rotor (68, 69), the combustor (122), the core turbine rotor (70, 71) and the power turbine rotor (67) from an inlet (54) into the flowpath (52) to an exhaust (56) from the flowpath (52).

An apparatus has: a first member; a shaft rotatable relative to the first member about an axis (A); and a seal system (100). The seal system has: a seal (102) carried by the first member and having a seal face (106); a seal carrier (150); a seat (104) carried by the shaft and having a seat face (108) in sliding sealing engagement with the seal face; and one or more springs (132) biasing the seal carrier relative to the first member so as to bias the seal face against the seat face. The seal carrier has: an axially-extending wall having an inner diameter (ID) surface (160); and a radially-extending wall having a first surface (158). The seal carrier axially-extending wall ID surface has a radially inwardly open groove (192) having a first sidewall and a second sidewall and a base. A wave-form split ring (190) contacts the first sidewall and biases the seal into engagement with the radial wall first surface.

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Zusammenfassung wird geladen …

Examples described herein provide a method that includes receiving engine data 208 from a sensor 206 associated with the engine 104. The method further includes associating a header with the engine data 208 to generate packaged engine data. The method further includes transmitting the packaged engine data to an aircraft communication unit 102, wherein the header provides for the aircraft communication unit 102 to transmit the packaged engine data to a ground station communication unit 106 via a communication protocol.

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A gas turbine engine includes a cracking device (74) that is configured to decompose an ammonia flow into a flow that contains more hydrogen (H2) than ammonia (NH3), a first separation device (76) that separates hydrogen downstream of the cracking device (74), wherein residual ammonia and nitrogen are exhausted as a residual flow (82). The separated flow contains more hydrogen than ammonia, and nitrogen is exhausted separately as a hydrogen flow (80). A combustor (46) is configured to receive and combust the hydrogen flow (80) from the separation device to generate a gas flow (58). A compressor section (44) is configured to supply compressed air to the combustor (46). A turbine section (48) is in flow communication with the gas flow (58) produced by the combustor (46) and is mechanically coupled to drive the compressor section (44).

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

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Zusammenfassung wird geladen …

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Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

A gas turbine engine includes a cracking device (70) that is configured to decompose a portion of an ammonia flow into a flow of component parts of the ammonia flow, a thermal transfer device that is configured to heat the ammonia flow to a temperature above 500 °C (932 °F), a combustor (46) that is configured to receive and combust the flow of component parts of the ammonia flow to generate a high energy gas flow (64), a compressor section (44) that is configured to supply compressed air to the combustor (46), and a turbine section (48) in flow communication with the high energy gas flow (64) produced by the combustor (46) and mechanically coupled to drive the compressor section (44).

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A maintenance scheduling system (300) for gas turbine engine components includes a computer system configured to receive a set of measured parameters for each gas turbine engine component in a plurality of substantially identical gas turbine engine components, and determine a variation model based on the set of measured parameters. Each of the gas turbine engine components is a single route component and has been utilized in a substantially identical single route. The computer system includes at least one simulated engine model. The simulated engine model is configured to determine a predicted operation of each gas turbine engine component in the plurality of substantially identical gas turbine engine components. A correlation system is configured to correlate variations in the set of parameters for each of the gas turbine engine components in the plurality of substantially identical gas turbine engine components with a set of the predicted operations of each gas turbine engine component in the plurality of substantially identical gas turbine engine components, thereby generating a predictive model based on the variations. A maintenance schedule module is configured determine a predictive model operable to determine a predicted wear pattern of a second single route component that is substantially identical to the gas turbine engine components.

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Zusammenfassung wird geladen …