A gas turbine engine component includes a passage and a geometrically segmented coating section adjacent the passage. The geometrically segmented coating section includes a wall that has a first side bordering the passage and a second side opposite the first side. The second side includes an array of cells, and there is a coating disposed over the array of cells. The coating defines an exterior side. A cooling passage extends through the wall and the coating. The cooling passage fluidly connects the passage and the exterior side.

An oil cooling system and method are provided for use with respect to a lubricated mechanical system within a bypass configured gas turbine engine. A surface cooler is fluidly linked to the lubricated mechanical system to receive oil from the lubricated mechanical system for cooling and reuse. In an embodiment, the surface cooler is mounted on an existing surface within the bypass airflow path of the bypass configured gas turbine engine to provide effective cooling while avoiding the introduction of additional aerodynamic disturbances in the bypass path. In an embodiment, the surface cooler is mounted on the fan casing or on a fan exit guide vane.

An oxide and non-oxide based ceramic matrix composite (CMC) component for a combustion turbine engine has a solidified ceramic core with a three-dimensional preform of ceramic fibers, embedded therein. Engineered surface features (ESFs) are cut into an outer surface of the core and fibers of the preform. A thermal barrier coat (TBC) is applied over and coupled to the core outer surface and the ESFs. The ESFs provide increased surface area and mechanically interlock the TBC, improving adhesion between the ceramic core and the TBC.

Turbine and compressor casing abradable component embodiments for turbine engines vary localized porosity or abradability through use of holes or dimple depressions of desired polygonal profiles that are formed into the surface of otherwise monolithic abradable surfaces or rib structures. Abradable porosity within a rib is varied locally by changing any one or more of hole/depression depth, diameter, array pitch density, and/or volume. In various embodiments, localized porosity increases and corresponding abradability increases axially from the upstream or forward axial end of the abradable surface to the downstream or aft end of the surface. In this way, the forward axial end of the abradable surface has less porosity to counter hot working gas erosion of the surface, while the more aft portions of the abradable surface accommodate blade cutting and incursion with lower likelihood of blade tip wear.

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.

A number of variations may include a method of optimizing inlet guide vane performance comprising: modifying an inlet guide vane to include at least one of a twist, a curve, a surface texture, a sealing feature, a tip leakage reduction feature, an airfoil having at least one component, or at least one channel.

An air guidance device for a turbomachine includes an air supply channel of a turbomachine engine. The supply channel has an upstream section and a downstream section connected together by a diverting section, the upstream section and the diverting section being connected together via an internal elbow and an external elbow. At the internal elbow, the internal surface has a groove extending longitudinally in the longitudinal direction of the supply channel and the longitudinal edges of which are widened in the direction of the downstream end of the upstream section of the supply channel.

A multilayer acoustic treatment panel including a first cellular-structure core sandwiched between a perforated skin and an intermediate skin; and a second cellular-structure core sandwiched between the intermediate skin and a continuous skin. The perforated skin includes at least one pair of high-porosity zones presenting a perforation ratio greater than a perforation ratio of a remainder of the perforated skin and including an inlet zone and an outlet zone longitudinally spaced apart from each other, the high-porosity zones of a given pair communicating through the first cellular-structure core and the intermediate skin with the two ends of a soundwave flow channel arranged in the second cellular-structure core.

A turbine nozzle disposed in a turbine includes a suction side extending between a leading edge of the nozzle and a trailing edge of the turbine nozzle in an axial direction and transverse to a longitudinal axis of the turbine nozzle, and extending a height of the nozzle in a radial direction along the longitudinal axis, a pressure side disposed opposite the suction side and extending between the leading edge of the turbine nozzle and the trailing edge of the turbine nozzle in the axial direction, and extending the height of the nozzle in the radial direction, and a bulge disposed on the suction side of the nozzle protruding relative to the other portion of the suction side in a direction transverse to a both the radial and axial directions.

A blade is used in fluid and includes: a base member; and a coat layer that is formed on the base member, a plurality of first grooves and a plurality of second grooves are formed on a surface of the coat layer, a pitch of the plurality of first grooves is different from a pitch of the plurality of second grooves.