RTX Corporation Patente

A rotating leaf spring seal (250; 350; 450) is disclosed. In various embodiments, the rotating leaf spring seal (250; 350; 450) includes a first annular ring (254; 454) configured for positioning against a support structure (245; 445); a hook section (272; 372; 472) having an outer surface (274) configured for sealing engagement with a tie shaft (220; 320; 420) and an inner surface (278) configured for receiving a hold down ring (280; 380); and a spring section (260; 360; 460) disposed between the first annular ring (254; 454) and the hook section (272; 372; 472).

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

An actively controlled squeeze film damper system (62) comprises a housing defining an annulus (68) receiving a damping fluid (84) during operation, a lubricant source supplying damping fluid (84) to the annulus (68), and a sensor assembly for measuring a parameter indicative of a compressibility of the damping fluid (84). A control device adjusts the compressibility of the damping fluid (84) within a predefined range.

A method for determining the presence of a particle while actively calculating and monitoring oil debris monitor phase angle in an oil system including collecting I and Q channel data from an oil debris monitor sensor; determining whether the I and Q data is symmetric; processing the I and Q channel data to identify a ferrous and nonferrous signal in response to the I and Q data being symmetric; processing the ferrous and nonferrous signals to determine if a particle is present; determining a symmetry factor from the I and Q channel data in response to the particle being present and confirming that the particle is present from the symmetry factor.

A repair 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 wherein each of the gas turbine engine components is an as-run component and has been exposed to a substantially identical general wear pattern. 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 repair module is configured to generate a repair formula based on the predictive model. The repair formula is configured to receive a set of measured parameters of an as-run gas turbine engine component and determine a suggested repair capable of repairing the as-run gas turbine engine component.

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

A dual-use air turbine system (40) for a gas turbine engine of an aircraft is provided. The dual-use air turbine system includes a variable area air turbine (42) mechanically linked to a spool of the gas turbine engine through a multi-speed gear set (44). The dual-use air turbine system also includes a plurality of valves (60; 62; 64; 66; 68) in pneumatic ducting (52) operable to direct an engine start air flow (54) through an inlet (58) of the variable area air turbine (42) and drive rotation of the spool during an engine start mode of operation. The valves are further operable to direct an engine bleed air flow (70) from a compressor section (24) of the gas turbine engine through the inlet of the variable area air turbine and drive rotation of the spool during an environmental control system active mode of operation.

An airfoil cluster (100) includes an inner diameter platform (102) and an outer diameter platform (104). A plurality of airfoils (47) extend radially from the inner diameter platform to the outer diameter platform. Each airfoil has a leading edge (106) and a trailing edge (108), and pressure and suction sides (110,112) between the leading and trailing edges. At least one of the airfoils includes a recessed portion (114) on the pressure side adjacent to the trailing edge.

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

A gas turbine engine (100) has a fan rotor (87) delivering air into a bypass duct defined between an outer fan case (94) and an outer interior housing. The fan rotor (87) also delivers air into a compressor section (77,79), a combustor (106), a turbine section (108H,108L). A chamber (97) is defined between the outer interior housing and an inner housing (113). The inner housing (113) contains the compressor section (77,79), the combustor (106) and the turbine section (108H,108L). A first conduit (110) taps hot compressed air to be cooled and passes the air to at least one heat exchanger (102). The air is cooled in the heat exchanger (102) and returned to a return conduit (112). The return conduit (112) passes the cooled air to at least one of the turbine section (108H,108L) and the compressor section (77,79). The heat exchanger (102) has a core exhaust plane (86). The turbine section (108H,108L) has at least a first and a downstream second rotor blade row (80,82), with the core exhaust plane (86) located downstream of a center plane (88) of the second blade row (82).

A method of fabricating a combustor thermal shield (108) comprising a combustor panel (110), a cooling feature (116), and an attachment feature (114), the combustor thermal shield (108) to be used in a gas turbine engine combustor (36), includes shaping a sheet of material used to form the combustor panel (110). The method also includes additively manufacturing the cooling feature (116) onto the sheet of material forming the combustor panel (110). The method also includes attaching the attachment feature (114) to the sheet of material forming the combustor panel (110). The method also includes curving the sheet of material forming the combustor panel (110) to achieve a curve profile according to a design of the gas turbine engine combustor (36).

Zusammenfassung wird geladen …

A method for forming an airfoil, a core for forming an airfoil, and an airfoil for a gas turbine engine are described. The method includes forming an airfoil body (302; 502; 992) about a core assembly (450), the core assembly comprising at least one core structure (452, 454, 456, 458; 781; 990) having a radially extending purge slot protrusion (462; 782; 994) extending from a portion of the at least one core structure, removing the core assembly from the airfoil body to form one or more internal cavities 530, 532, 534, 536; 632, 634; 783), wherein at least one internal cavity has a cavity extension (642; 784) defined by the purge slot protrusion (462; 782; 994), and forming a squealer pocket (320; 620; 786) in a tip of the formed airfoil body, wherein a cavity purge slot (326; 526; 628; 790) is formed at the cavity extension to fluidly connect the respective internal cavity with the squealer pocket. The radially extending protrusion has a radial height that is equal to or less than five times a width thereof.

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

A combustor of a gas turbine engine including a first forward liner panel (72A-1, 74A-1). The forward liner panel including a liner panel forward edge (130), a liner panel aft edge (132) opposite of the liner panel forward edge, and a first liner panel side edge (134) between the liner panel aft edge and the liner panel forward edge. The first liner panel side edge is non-perpendicular to the liner panel forward edge and oriented at an angle about perpendicular to a swirl flow (I, O) in the combustor.

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

A gas turbine engine (20) includes a turbine section (28) configured to rotate about an engine axis (X) relative to an engine static structure (36). A fan section (22) is configured to rotate about the engine axis (X) relative to the engine static structure (36). A geared architecture (48) operatively connects the fan section (22) to the turbine section (28). The geared architecture (48) includes an input gear (60; 160) and an output gear (62; 162) both configured to rotate about the engine axis (X). A set of idler gears (66; 166, 266) are arranged radially outward relative to and operatively coupling with the input and output gears (60, 62; 160, 162).

Zusammenfassung wird geladen …

Zusammenfassung wird geladen …

A low profile un-lensed non-intrusive stress measurement system probe (200) may comprise a housing (300) comprising a first channel (510) and an optical face (506), a first hypotube (310) disposed within the first channel (510) and coupled at a sensing aperture (508) in the optical face (506), and a plurality of optical fibers (412) disposed within the first hypotube (310), wherein the first hypotube (310) executes a bend between 45° and 90° within the housing (300).

Zusammenfassung wird geladen …