Abstract: Clevis-type pin attachment mounts for ceramic-matric-composite (CMC) blades (50) accommodate varying thermal expansion rates between ceramic blades and the mating engine rotor disc (46). A two-dimensional array of apertures (124, 126, 128, and 130) the CMC blade shank (70) receives of rows of load-carrying pins (132, 134, 136, and 138). Tensile loads applied to the pin and aperture array are distributed within the blade shank, so that applied tensile load stress is split between successive rows of apertures and pins, so that each row of apertures carries its own tensile load plus aggregate tensile load of all other rows of apertures that are closer to the blade tip. Axial gaps (GA) between tips of load-carrying pins and partial-depth apertures in clevis attachment pieces (100, 102) provide compressive loading on the blade shank (70).

Abstract: A component for a combustion turbine engine has a ceramic matrix composite (“CMC”) central substrate with a substrate rib that bridges opposed substrate sidewalls. The central substrate includes an embedded first pattern of reinforcing fibers, which is laid-up from a pair of fabric reinforcement sheets. The respective sheets include a spine with coplanar, flanking strips. Alternate rows of strips on each sheet are folded into projecting pleats. The respective sheets are oriented with their spines in opposed and mutually spaced relationship. Spines of the respective sheets are embedded within their respective substrate sidewalls. Pleat rows of the respective sheets are embedded within the substrate rib and its opposing respective first or second sidewall.

Abstract: A blade and tip shroud assembly for a gas turbine engine. The assembly includes at least one blade with an airfoil extending span-wise along a radial direction. Each blade includes a blade tip. At least one tenon extends radially out from the blade tip. A ceramic matrix composite (CMC) tip shroud is positioned along the blade tip. The CMC tip shroud extends along a circumferential direction and includes an upstream edge and a downstream edge spaced apart from each other in an axial direction. The CMC tip shroud includes a radially inner diameter (ID) surface adjoining the blade tip and a radially outer diameter (OD) surface opposite to the radially ID surface. A securing mechanism is connected to the OD surface of the CMC tip shroud and attached to the at least one tenon that extends through the CMC tip shroud.

Abstract: A gas turbine engine blade (10), including a base portion (12) having a cast wall, and a tip portion (14) attached to the base portion and having a wall (60) formed by an additive manufacturing process. The tip portion wall may be formed to be solid and less than 2 mm in thickness, or it may be corrugated and be greater than 2 mm in thickness. Openings (80) defining the wall corrugations may be semi-circular, rectangular, trapezoidal, or elliptical in cross-sectional shape. The resulting blade has lower tip mass while retaining adequate mechanical properties. The tip portion may be formed to have a directionally solidified grain structure on a base portion having an equiaxed grain structure.

Abstract: A ceramic matrix composite (“CMC”) component, such as a turbine blade for a combustion turbine engine that has a fiber-reinforced, solidified ceramic substrate. The substrate has an inner layer of fibers, for enhancing structural strength of the component. An outer layer of fibers defines voids therein. A thermal barrier coat (“TBC”) is applied over and coupled to the outer layer fibers, filling the voids. The voids provide increased surface area and mechanically interlock the TBC, improving adhesion between the fiber-reinforced ceramic substrate and the TBC.

Abstract: A turbine arrangement including a rotor and a stator surrounding the rotor and comprising guide vane segments, each guide vane segment comprising an airfoil and a radially inner vane platform. A seal arrangement includes a static seal inward from the inner vane platforms and having a radially extending face plate, first and second cylindrical seal walls extending from outer and inner ends of the annular face plate, an annular seal plate extending radially from the second cylindrical seal wall, and an angel wing extending between the first cylindrical seal wall and the annular seal plate to define a first annular cavity and a second annular cavity. Circumferentially spaced cut-outs define passages through the annular seal plate between the first and second annular cavities and are aligned with fasteners that attach the annular face plate to a support ring for supporting the inner vane platform.

Abstract: A modular turbine blade assembly (10) usable in a gas turbine engine (12) and formed from an airfoil (28) and an independent, modular platform (16) supported by one or more clevis arm supports (14) extending radially inward from the modular platform (16) to a disk is disclosed. The clevis arm support may support the modular platform while a separate dovetail attachment supports the generally hollow airfoil. The clevis arm support (14) may be formed from at least two arms (20, 22) designed to reduce stress from a pin receiving orifice (24) at a distal end (26) of the two arms (20, 22) to the platform (16). The independent arms (20, 22) minimize stress concentrations caused by centrifugal loading in the support. The arms (20, 22) may be modified independently of each other, such as thickness and support angle. The clevis arm support (14) enables use of a modular platform system for the modular turbine blade (10).

Abstract: A gas turbine engine blade (10), including a base portion (12) having a cast wall, and a tip portion (14) attached to the base portion and having a wall (60) formed by an additive manufacturing process. The tip portion wall may be formed to be solid and less than 2 mm in thickness, or it may be corrugated and be greater than 2 mm in thickness. Openings (80) defining the wall corrugations may be semi-circular, rectangular, trapezoidal, or elliptical in cross-sectional shape. The resulting blade has lower tip mass while retaining adequate mechanical properties. The tip portion may be formed to have a directionally solidified grain structure on a base portion having an equiaxed grain structure.

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: A vane array adapted to be coupled to a vane carrier within a gas turbine engine is provided comprising: a plurality of elongated airfoils comprising at least a first airfoil and a second airfoil located adjacent to one another; a U-ring; first connector structure for coupling a radially inner end section of each of the first and second airfoils to the U-ring; second connector structure for coupling a radially outer end section of each of the first and second airfoils to the vane carrier; a platform extending between the first and second airfoils; and platform connector structure for coupling the platform to one of the U-ring and the vane carrier.

Abstract: A vane array adapted to be coupled to a vane carrier within a gas turbine engine is provided comprising: a plurality of elongated airfoils comprising at least a first airfoil and a second airfoil located adjacent to one another; a U-ring; first connector structure for coupling a radially inner end section of each of the first and second airfoils to the U-ring; second connector structure for coupling a radially outer end section of each of the first and second airfoils to the vane carrier; a platform extending between the first and second airfoils; and platform connector structure for coupling the platform to one of the U-ring and the vane carrier.

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.