Gas turbine engines and methods associated therewith

Abstract

A method of forming a gas turbine engine component, the method including forming a plurality of cooling apertures in a preform structure of the component, the plurality of cooling apertures of the preform structure comprising a first cooling aperture and a second cooling aperture, wherein cross-sectional shapes of the first and second cooling apertures of the preform structure are different from one another, as measured in a same relative plane; and applying a coating to at least a portion of the preform structure to form the component, wherein a cross-sectional shape of the first and second cooling apertures of the component are approximately the same as one another, as measured in the same relative plane.

What is claimed is: 
     
         1 . A gas turbine engine component, comprising:
 a preform structure defining a plurality of cooling apertures, the plurality of cooling apertures including a first cooling aperture defining a centerline axis; and   a coating disposed over at least a portion of the preform structure; and   wherein the first cooling aperture includes a sloped landing for receiving the coating, the sloped landing defining a first slope profile along a length of the first cooling aperture and offset from the centerline axis, wherein the coating is disposed over the sloped landing and defines a second slope profile along the length of the first cooling aperture, and wherein the first slope profile and the second slope profile are different from one another.   
     
     
         2 . The gas turbine engine component of  claim 1 , wherein the coating comprises at least one of a thermal barrier coating (TBC), a bond coating, or an environmental barrier coating. 
     
     
         3 . The preform structure of  claim 1 , wherein the first cooling aperture is an additively manufactured cooling hole aperture. 
     
     
         4 . The gas turbine engine component of  claim 1 , wherein the coating has an uneven thickness. 
     
     
         5 . The gas turbine engine component of  claim 4 , wherein the uneven thickness corresponds with a dispersion pattern used in applying the coating. 
     
     
         6 . The gas turbine engine component of  claim 1 , wherein the coating is a plasma deposition coating or a direction coating. 
     
     
         7 . The gas turbine engine component of  claim 6 , wherein directional coating is performed in a direction within a range between 25 degrees and 90 degrees with respect to a relative plane. 
     
     
         8 . The gas turbine engine component of  claim 1 , wherein the preform structure forms a portion of a blade. 
     
     
         9 . The gas turbine engine component of  claim 1 , wherein the sloped landing is angularly offset from the centerline axis by at least 1 degree. 
     
     
         10 . The gas turbine engine component of  claim 1 , wherein the sloped landing is angularly offset from the centerline axis by at least 10 degrees. 
     
     
         11 . A gas turbine engine component, comprising:
 a preform structure defining a plurality of cooling apertures, the plurality of cooling apertures including a first cooling aperture; and   a coating disposed over at least a portion of the preform structure; and   wherein the first cooling aperture includes an inlet bore, an outlet, and a through passage therebetween, and wherein the outlet includes a landing, sloped away from the through passage, the landing configured for receiving the coating, the landing defining a first slope profile along a length of the first cooling aperture, wherein the coating is disposed over the landing and defines a second slope profile along the length of the first cooling aperture, and wherein the first slope profile and the second slope profile are different from one another.   
     
     
         12 . The gas turbine engine component of  claim 11 , wherein the coating comprises at least one of a thermal barrier coating (TBC), a bond coating, or an environmental barrier coating. 
     
     
         13 . The preform structure of  claim 11 , wherein the first cooling aperture is an additively manufactured cooling hole aperture. 
     
     
         14 . The gas turbine engine component of  claim 11 , wherein the coating has an uneven thickness. 
     
     
         15 . The gas turbine engine component of  claim 14 , wherein the uneven thickness corresponds with a dispersion pattern used in applying the coating. 
     
     
         16 . The gas turbine engine component of  claim 11 , wherein the coating is a plasma deposition coating or a direction coating. 
     
     
         17 . The gas turbine engine component of  claim 16 , wherein directional coating is performed in a direction within a range between 25 degrees and 90 degrees with respect to a relative plane. 
     
     
         18 . The gas turbine engine component of  claim 11 , wherein the preform structure forms a portion of a blade. 
     
     
         19 . The gas turbine engine component of  claim 11 , wherein the landing includes a portion of the through passage, the portion of the through passage angularly offset from a centerline axis of the inlet bore. 
     
     
         20 . The gas turbine engine component of  claim 19 , wherein the landing is angularly offset from the centerline axis by at least 1 degree.