A turbomachine includes a shroud and a hub spaced apart from the shroud to channel an airflow along a direction. The turbomachine includes a plurality of airfoils coupled between the shroud and the hub. At least one airfoil of the plurality of airfoils includes a leading edge spaced apart from a trailing edge in the direction of the airflow and a pressure side opposite a suction side. The at least one airfoil has a chord that is 0% chord at the leading edge and 100% chord at the trailing edge, and a span defined as 0% at the hub and 100% at the shroud. The turbomachine includes at least one local vortex generator array defined onto the hub between 0% chord to ?10% chord at 0% span so as to extend forward of the leading edge.

Turbine and compressor casing/housing abradable component embodiments for turbine engines, have abradable surfaces with ridges projecting from the abradable surface, separated by grooves. The ridges have one or both sidewalls inclined against the opposing turbine blade tip rotational direction for redirecting and/or dissipating blade tip gap leakage airflow energy. In some embodiments the ridge tip and/or groove base have inclined profiles for redirecting airflow leakage away from the blade tip gap. In some embodiments, the inclined ridge tip profile provides a progressive wear zone that increases abradable surface area as the inclined ridge is abraded by the rotating blade tip.

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.

Turbine and compressor casing abradable components for turbine engines include abradable surfaces with a zonal system of forward (zone A) and rear or aft sections (zone B) surface features. The zone A surface profile comprises an array pattern of non-directional depression dimples, or upwardly projecting dimples, or both, in the abradable surface. The dimpled forward zone A surface features reduce surface solidity in a controlled manner, to help increase abradability during blade tip rubbing incidents, yet they provide sufficient material to resist incoming hot working fluid erosion of the abradable surface. In addition, the dimples provide generic forward section aerodynamic profiling to the abradable surface, compatible with different blade airfoil-camber profiles. The aft zone B surface features comprise an array pattern of ridges and grooves.

Turbine and compressor casing/housing abradable component embodiments for turbine engines, have abradable surfaces with asymmetric forward and aft ridge surface area density. The forward ridges have greater surface area density than the aft ridges to compensate for greater ridge erosion in the forward zone during engine operation and reduce blade tip wear in the aft zone. Some abradable component embodiments increase forward zone ridge surface area density by incorporating wider ridges than those in the aft zone.

Delamination of thermal barrier coatings (TBC's) (276) from superalloy substrates (262) of components (260) for turbine engines (80), such as engine blades (92), vanes (104, 106), or castings in transitions (85), is inhibited during subsequent cooling passage (270) formation. Partially completed cooling passages (264), which have skewed passage paths that end at a terminus (268), which is laterally offset from the passage entrance (266), are formed in the superalloy component (260) prior to application of the TBC layer(s) (276). The skewed, laterally offset path of each partially completed cooling passage (264) establishes an overhanging shield layer (269) of superalloy material that protects the TBC layer (276) during completion of the cooling passage (270).

Cooling passages (99, 105) are formed in components for combustion turbine engines, such as blades (92), vanes (104, 106), ring segments (110) or castings in transitions (85), during investment casting, through use of ceramic shell inserts (130) within the casting mold (152). Ceramic posts (134) formed in the ceramic shell insert (130) have profiles conforming to corresponding profiles of partially completed cooling passages (156). Posts (134) are removed after superalloy component casting, forming the partially completed cooling passages, which are subsequently completed by removing remaining superalloy material along the cooling passage path.

An auxetic (NPR) structure includes a plurality of vertical intersecting dimpled sheets, each dimpled sheet exhibiting a negative Poisson's ratio, each dimpled sheet intersects two adjacent dimpled sheets creating a rectangular tubular structure, and having a portion of each dimpled sheet projecting outward from its intersection with an adjacent dimpled sheet, the amplitude of each dimple on the plurality of dimpled sheets is such that no overhanging surface of the dimpled sheet exceeds an angle threshold for printability without support structures.

A gas turbine engine includes a fan section, a compressor section, a combustion section, and a turbine section in serial flow arrangement, and defining an engine centerline extending between a forward direction and an aft direction. A disk includes a slot for mounting a composite airfoil to the disk. An axial retainer couples to the disk and secures the composite airfoil to the disk. A compliant portion positioned at the composite airfoil abuts the composite airfoil during operation of the gas turbine engine to secure the composite airfoil to the disk.

A gas turbine engine includes a fan delivering air into a bypass duct defined between a nacelle and an inner core housing. The inner core housing receives a compressor section, a turbine section and a combustor. The nacelle has an inner periphery with a forwardmost point, and receives an acoustic structure on the inner periphery adjacent the forwardmost point. The acoustic structure is defined by a three dimensional array of interconnected resonators, with the interconnected resonators extending in a radial direction, a circumferential direction and an axial direction all defined about a rotational axis of the engine, with the interconnected resonators having a larger cross-sectional area central body, and six members connecting the central body of the resonators to adjacent resonators at respective central bodies. A perforated face sheet is inward of the three dimensional array of interconnected resonator and an anti-icing system.