Abstract: A method of forming a ceramic matrix composite (CMC) vane comprises drawing a continuous fiber sheet over at least one contoured surface to form a single-folded sheet. The single-folded sheet comprises an airfoil portion, a first platform at a first end of the airfoil portion, and an integral transition region between the airfoil portion and the first platform. The method further comprises mounting the airfoil portion on a mandrel, attaching a second platform to a second, opposing end of the airfoil portion to form a vane preform, and densifying the vane preform.

Abstract: A testing arrangement includes a test specimen that is formed of a fiber composite that has first, second, and third groups of fiber ply layers. Each fiber ply layer is curved so as to have an elbow. The elbows are joined back-to-back at a central region to form first, second, and third arms that extend outwardly from the central region. The elbows define a space in the central region, and there is a space-filler within the space. The test specimen is mounted in a testing fixture such that the first and second arms are in four-point bend loading and the third arm is oriented vertically. The testing fixture is actuatable to apply an actuated load. There is an applied load on the third arm. The applied load and the actuated load cause stress in the space-filler for evaluating mechanical behavior of the space-filler.

Abstract: An airfoil includes an airfoil wall that has a wishbone-shaped fiber layup structure. The structure includes two arms that merge into a single leg. The single leg includes fiber tows from each of the arms that are interwoven. The single leg forms at least a portion of the trailing edge of the airfoil wall. Each of the arms includes a first section that is distal from the single leg and a second section that is proximal to the single leg. The first section defines a first thickness, the second section defines a second thickness that is greater than the first thickness, and the single leg defines a third thickness that is greater than the second thickness.

Abstract: A method for processing a CMC airfoil includes nesting an airfoil fiber preform in a cavity of a fixture that has first and second tool segments, closing the fixture by rotating a first tool segment about a hinge, the closing causing the tool segments to clamp on a tail portion of the fiber preform and thereby conform the tail portion to the fixture. While in the fixture, the fiber preform is then partially densified with an interface coating material to form a partially densified fiber preform. While still in the fixture, one or more cooling holes are drilled into the trailing edge of the partially densified fiber preform. After the drilling, the partially densified fiber preform is removed from the fixture and further densified with a ceramic matrix material to form a fully densified CMC airfoil.

Abstract: A method of forming a ceramic matrix composite (CMC) vane comprises drawing a continuous fiber sheet over at least one contoured surface to form a single-folded sheet. The single-folded sheet comprises an airfoil portion, a first platform at a first end of the airfoil portion, and an integral transition region between the airfoil portion and the first platform. The method further comprises mounting the airfoil portion on a mandrel, attaching a second platform to a second, opposing end of the airfoil portion to form a vane preform, and densifying the vane preform.

Abstract: A multiple static vane component includes a plurality of airfoils each formed of ceramic matrix composite materials. Each of the airfoils are attached to an inner platform and an outer platform both formed of ceramic matrix composite materials. There is a plurality of individual parts forming the plurality of airfoils, the inner platform or the outer platform, bonded to each other with a braze joint. A gas turbine engine and a method are also disclosed.

Abstract: A method for processing a CMC airfoil includes nesting an airfoil fiber preform in a cavity of a fixture that has first and second tool segments, closing the fixture by rotating a first tool segment about a hinge, the closing causing the tool segments to clamp on a tail portion of the fiber preform and thereby conform the tail portion to the fixture. While in the fixture, the fiber preform is then partially densified with an interface coating material to form a partially densified fiber preform. While still in the fixture, one or more cooling holes are drilled into the trailing edge of the partially densified fiber preform. After the drilling, the partially densified fiber preform is removed from the fixture and further densified with a ceramic matrix material to form a fully densified CMC airfoil.

Abstract: An airfoil includes an airfoil wall that has a wishbone-shaped fiber layup structure. The structure includes two arms that merge into a single leg. The single leg includes fiber tows from each of the arms that are interwoven. The single leg forms at least a portion of the trailing edge of the airfoil wall. Each of the arms includes a first section that is distal from the single leg and a second section that is proximal to the single leg. The first section defines a first thickness, the second section defines a second thickness that is greater than the first thickness, and the single leg defines a third thickness that is greater than the second thickness.

Abstract: A method for processing a CMC airfoil includes nesting an airfoil fiber preform in a cavity of a fixture that has first and second tool segments, closing the fixture by rotating a first tool segment about a hinge, the closing causing the tool segments to clamp on a tail portion of the fiber preform and thereby conform the tail portion to the fixture. While in the fixture, the fiber preform is then partially densified with an interface coating material to form a partially densified fiber preform. While still in the fixture, one or more cooling holes are drilled into the trailing edge of the partially densified fiber preform. After the drilling, the partially densified fiber preform is removed from the fixture and further densified with a ceramic matrix material to form a fully densified CMC airfoil.

Abstract: A gas turbine engine includes a ceramic matrix composite (CMC) vane arc segments that are arranged in a circumferential row. Each of the CMC vane arc segments includes an airfoil section that defines first and second side walls, leading and trailing ends, and first and second radial ends. At the first radial end, the airfoil section has a single-sided platform that and the second side wall has a bearing surface. The single-sided platform of each of the CMC vane arc segments in the circumferential row is situated to bear against the bearing surface of the next of the CMC vane arc segments in the circumferential row.

Abstract: An airfoil vane multiplet includes airfoil tubes that each have airfoil tube fiber plies. A plurality of the airfoil tube fiber plies extend along the airfoil tube and turn to project outwardly from the airfoil tube. Connector fiber plies include airfoil holes through which the airfoil tubes extend. The connector fiber plies are interleaved with the plurality of airfoil tube fiber plies such that the airfoil tubes are secured together.

Abstract: A vane multiplet includes a plurality of airfoils that each have a flared end, a common platform piece, a plurality of seal, and a plurality of retainers. The common platform has airfoil sockets that each define an airfoil opening that is circumscribed by a groove. The flared ends of the airfoils are seated in the airfoil sockets such that the grooves and the airfoils together form seal channels that have an open side. The seals are disposed in the seal channels. The retainers have airfoil-shaped profiles and are disposed in the airfoil sockets to bound the open sides of the seal channels and retain the seal in the seal channel.

Abstract: A gas turbine engine includes a turbine section that has a plurality of turbine vanes. Each of the turbine vanes includes inner and outer platforms and an airfoil section that extends there between. The airfoil section is hollow and rib-less and has a first end at the outer platform and a second end at the inner platform. The airfoil section is tangentially bowed from the first end to the second end with a radius of curvature that is from 17 centimeters to 130 centimeters.

Abstract: A gas turbine engine includes a turbine section that has a plurality of turbine vanes. Each of the turbine vanes includes inner and outer platforms and an airfoil section that extends there between. The airfoil section is hollow and rib-less and has a first end at the outer platform and a second end at the inner platform. The airfoil section is tangentially bowed from the first end to the second end with a radius of curvature that is from 17 centimeters to 130 centimeters.

Abstract: A turbine airfoil usable in a turbine engine and having at least one ambient air cooling system is disclosed. At least a portion of the cooling system may include one or more cooling channels configured to receive ambient air at about atmospheric pressure. The ambient air cooling system may have a tip static pressure to ambient pressure ratio of at least 0.5, and in at least one embodiment, may include a tip static pressure to ambient pressure ratio of between about 0.5 and about 3.0. The cooling system may also be configured such that an under root slot chamber in the root is large to minimize supply air velocity. One or more cooling channels of the ambient air cooling system may terminate at an outlet at the tip such that the outlet is aligned with inner surfaces forming the at least one cooling channel in the airfoil to facilitate high mass flow.

Abstract: A turbine blade is provided comprising: a root; an airfoil comprising an external wall extending radially from the root and having a radially outermost portion; and a damping structure. The external wall may comprise first and second side walls joined together to define an inner cavity of the airfoil. The damping structure may be positioned within the airfoil inner cavity and coupled to the airfoil so as to define a tuned mass damper.

Abstract: An outer rim seal arrangement (10), including: an annular rim (70) centered about a longitudinal axis (30) of a rotor disc (31), extending fore and having a fore-end (72), an outward-facing surface (74), and an inward-facing surface (76); a lower angel wing (62) extending aft from a base of a turbine blade (22) and having an aft end (64) disposed radially inward of the rim inward-facing surface to define a lower angel wing seal gap (80); an upper angel wing (66) extending aft from the turbine blade base and having an aft end (68) disposed radially outward of the rim outward-facing surface to define a upper angel wing seal gap (80, 82); and guide vanes (100) disposed on the rim inward-facing surface in the lower angel wing seal gap. Pumping fins (102) may be disposed on the upper angel wing seal aft end in the upper angel wing seal gap.

Abstract: A cooling channel (36, 36B, 63-66) cools inner surfaces (48, 50) of exterior walls (41, 43) of a component (20, 60). Interior side surfaces (52, 54) of the channel converge to a waist (W2), forming an hourglass shaped transverse profile (46). The inner surfaces (48, 50) may have fins (44) aligned with the coolant flow (22). The fins may have a transverse profile (56A, 56B) highest at mid-width of the inner surfaces (48, 50). Turbulators (92) may be provided on the side surfaces (52, 54) of the channel, and may urge the coolant flow toward the inner surfaces (48, 50). Each turbulator (92) may have a peak (97) that defines the waist of the cooling channel. Each turbulator may have a convex upstream side (93). These elements increase coolant flow in the corners (C) of the channel to more uniformly and efficiently cool the exterior walls (41, 43).

Abstract: Platforms (36, 38) span between turbine blades (23, 24, 25) on a disk (32). Each platform may be individually mounted to the disk by a pin attachment (42). Each platform (36) may have a rotationally rearward edge portion (50) that underlies a forward portion (45) of the adjacent platform (38). This limits centrifugal bending of the rearward portion of the platform, and provides coolant sealing. The rotationally forward edge (44A, 44B) of the platform overlies a seal element (51) on the pressure side (28) of the forwardly adjacent blade, and does not underlie a shelf on that blade. The pin attachment allows radial mounting of each platform onto the disk via tilting (60) of the platform during mounting to provide mounting clearance for the rotationally rearward edge portion (50). This facilitates quick platform replacement without blade removal.

Abstract: An outer rim seal arrangement (10), including: an annular rim (70) centered about a longitudinal axis (30) of a rotor disc (31), extending fore and having a fore-end (72), an outward-facing surface (74), and an inward-facing surface (76); a lower angel wing (62) extending aft from a base of a turbine blade (22) and having an aft end (64) disposed radially inward of the rim inward-facing surface to define a lower angel wing seal gap (80); an upper angel wing (66) extending aft from the turbine blade base and having an aft end (68) disposed radially outward of the rim outward-facing surface to define a upper angel wing seal gap (80, 82); and guide vanes (100) disposed on the rim inward-facing surface in the lower angel wing seal gap. Pumping fins (102) may be disposed on the upper angel wing seal aft end in the upper angel wing seal gap.