RTX Corporation Patente

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A manufacturing method is provided. During this method, a preform component (60') for a turbine engine is provided that includes a substrate (74). A meter section (102) of a cooling aperture (64) is formed in the substrate (74'). An external coating (76',78') is applied over the substrate. At least a portion of the substrate (74') and the external coating (76',78') is scanned with an imaging system (156) to provide scan data indicative of an internal structure of the portion of the substrate (74') and the external coating (76',78'). A diffuser section of the cooling aperture is formed in the external coating (76',78') and the substrate (74') based on the scan data.

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An airfoil section (66) has an airfoil wall (68) that defines a leading end (68a), a trailing end (68b), and first and second sides (68c, 68d) that join the leading end (68a) and the trailing end (68b). The airfoil wall (68) circumscribes an internal core cavity (70). A skincore passage (74) is embedded in the first side (68c) of the airfoil wall (68) between inner and outer portions (76, 78) of the first side (68c) of airfoil wall (68). The skincore passage (74) defines a passage forward edge (74a) and a passage aft edge (74b). The first side (68c) of the airfoil wall (68) has a portion (86) that tapers in thickness at the passage aft edge (74b). A cooling passage (80) is embedded in the second side (68d) of the airfoil wall (68) between inner and outer portions (82, 84) of the second side (68d) of airfoil wall (68). The inner portion (82) of the second side (68d) of the airfoil wall (68) includes a protrusion (88) located across the internal core cavity (70) from the portion (86) of the first side (68c) that tapers in thickness.

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A brush seal system includes a component including a first mating surface (136). The brush seal system further includes a brush seal (110) including a brush seal backing plate (112), a retaining ring (114), and a plurality of bristles (116) retained between the brush seal backing plate (112) and the retaining ring (114). The brush seal backing plate (112) includes a second mating surface (138) mounted to the first mating surface (136). The brush seal system further includes at least one cooling channel (128) extending from an exterior side (130) of the component to an interior side (132) of the component so as to bypass the brush seal (110).

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A seal (100) includes a seal body (102) disposed about a seal axis (104) and configured to be mounted to a first component (60) at a second axial end (110) of the seal body (102). The seal body (102) includes an interior surface defining a seal gland (116) circumferentially extending about the seal axis (104). A packing (120) is disposed within the seal gland (116). A retaining ring (132) is in communication with a second radial side of the packing (120) and is disposed within the seal gland (116). A plurality of exciter springs (140) are mounted to a second radial side of the retaining ring (132). The plurality of exciter springs (140) are biased against the interior surface of the seal body (102) and are configured to center the packing (120) within the seal body (102). The packing (120) is configured to receive a second component (52) and form a seal interface between the packing (120) and the second component (52). The seal body (102) is configured to form a portion of a passage (106).

Combustor mounting structures (400; 500; 600; 700; 800; 900; 1000; 1100; 1200) for gas turbine engines (300) are described. The combustor mounting structures (400; 500; 600; 700; 800; 900; 1000; 1100; 1200) include a shell portion (326; 402; 502; 602; 702; 802; 902; 1002; 1102; 1202) defining an outer ring arranged at an outer radius relative to a central axis, a combustor connection element (320; 404; 508; 608; 708; 808; 908; 1008; 1108; 1208) defining an inner ring arranged at an inner radius that is less than the outer radius relative to the central axis, and a plurality of struts (328; 406; 506; 606; 706; 806; 906; 1006; 1106; 1206) extending radially between and connecting the shell portion (326; 402; 502; 602; 702; 802; 902; 1002; 1102; 1202) to the combustor connection element (320; 404; 508; 608; 708; 808; 908; 1008; 1108; 1208), wherein one or more flow apertures (410) are defined between the shell portion (326; 402; 502; 602; 702; 802; 902; 1002; 1102; 1202) and the combustor connection element (320; 404; 508; 608; 708; 808; 908; 1008; 1108; 1208) in a radial direction and between adjacent struts of the plurality of struts (328; 406; 506; 606; 706; 806; 906; 1006; 1106; 1206) in a circumferential direction.

A blade outer air seal assembly (104) includes a support structure (110). A blade outer air seal (106) has a plurality of seal segments (105) that extend circumferentially about an axis and mounted in the support structure via a carrier (112). The carrier (112) has a plurality of carrier segments that extend circumferentially about the axis. At least one of the seal segments (105) have a base portion (124) that extends between a first circumferential side (C1) and a second circumferential side (C2) and from a first axial side (A1) to a second axial side (A2). A first wall (120, 140) axially spaced from a second wall (121, 141). The first and second walls (120, 121, 140, 141) extend from the base portion (124) to a first outer portion to form a first passage (138). The first wall (120, 140) has at least one slot (130) engaged with a first carrier hook (141) on one of the plurality of carrier segments. At least one of the carrier segments have a carrier window (114) engaged with a support structure hook (160, 162) on the support structure (110).

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An apparatus for dislodging material (400) within an airfoil element is disclosed. In various embodiments, the apparatus (400) includes a robotic arm (402) configured for multi-axis movement of a mounting head (406) with respect to the airfoil element; a vibration mechanism (408) connected to the mounting head (406); and a tip member (412) connected to the vibration mechanism (408) and configured for vibratory contact with a surface of the airfoil element.

A turbofan engine (20) includes a fan section (22) including a plurality of fan blades (42), a first spool including a first turbine section, a second spool including a second turbine section, a first electric machine (62), and a second electric machine (64). A gear system (48, 48') is coupled between the first spool, the second spool, the first electric machine (62), the second electric machine (64) and the fan section (22) for driving the fan section (22).

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A combustor (108) for a gas turbine engine (100; 200) includes a combustor liner (100) and a dilution chute (114; 214) integral and conformal with the combustor liner (110) to provide an outlet into the combustor (108) for fuel, wherein the dilution chute (114; 214) has at least one wall (122; 222) with a tapered edge (118; 218) extending into an interior of the combustor (108).

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A blade outer air seal segment 130 including a radially outward surface 136, a radially inward surface 134 oriented away from the radially outward surface 136, and a cooling channel 180 located between the radially outward surface 136 and the radially inward surface 134. The blade outer air seal segment 130 also including a stress-relief boss 162 extending into the cooling channel 180 and an inlet orifice 170 fluidly coupled to the cooling channel 180 through the stress-relief boss 162.

Rotor systems (400) including an engine shaft (408), a forward hub (410), a rear hub (412), a rotor disk (402c) arranged between the forward hub (410) and the rear hub (412), and a seal tube (424) configured to define a forward hub compartment (420) and a rear hub compartment (422). The forward hub compartment (420) is defined forward of the rotor disk (402c) and the rear hub compartment (422) is defined aft of the rotor disk (402c). The seal tube (424) is connected at a forward end (428) to at least one of the rotor disk (402c) and the engine shaft (408) and at a rear end (432) to at least one of the rear hub (412) and the engine shaft (408) and the seal tube (424) includes at least one axial compliance element (426) configured to enable axial extension and compression of the seal tube (424) in an axial direction along the engine shaft (408).

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A method for repairing a blade (20)wherein the blade comprises a metallic substrate (22) shaped to define an airfoil having a tip. A coating (34) is on the tip. The method comprises: machining to at least partially remove the coating (34); plating a nickel-based base layer (32'); and plating an abrasive layer (34') comprising a nickel-based matrix (36) and an abrasive (38).

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Component for gas turbine engines (20) are described. The components include an airfoil body (402; 502) having leading and trailing edges (404, 406) and pressure and suction sides (408, 410). The airfoil has a leading edge cavity (412; 512) located proximate the leading edge (404) defined between the leading edge (404) and a separator rib (426; 526) and between the pressure side (408) and the suction side (410). An insert member (418; 518; 618) is installed within the leading edge cavity (412; 512). The insert member (418; 518; 618) has one or more metering flow apertures (420; 520; 620) at an aft end and one or more impingement apertures (422; 522; 622) at a forward end and at least one axially extending rib (530; 630) along an exterior surface thereof. At least one axial extending flow channel passage (532; 632) is defined along the axial extending rib (530; 630) between an exterior of the insert member (418; 518; 618) and an interior of the airfoil body (402; 502). Air flow through the metering flow apertures (420; 520; 620) flows into the axial extending flow channel passage (532; 632) and flows forward toward the leading edge (404).

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A gas turbine engine includes a fan rotor driven by a fan drive turbine about an axis through a gear reduction to reduce a speed of the fan rotor relative to a speed of the fan drive turbine. A fan case (108) surrounds the fan rotor, and a core engine with a compressor section, including a low pressure compressor. The fan rotor delivers air into a bypass duct defined between the fan case (108) and the core engine. A rigid connection (124) between the fan case (108) and the core engine includes a plurality of H-frame (116) connecting members rigidly connected to the fan case (108), and to the core engine. The H-frame (116) connecting members each are defined by two rigid legs (118) which extend between the fan case (108) and to the core engine, along directions which are generally parallel to each other. A plurality of non-structural fan exit guide vanes are provided with an acoustic feature to reduce noise and are rigidly mounted to at least one of the fan case (108) and the core engine.