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.

A method of infiltrating a fiber preform comprises positioning an assembly in a process chamber, where the assembly includes a tool comprising through-holes, a fiber preform constrained within the tool, and a sacrificial preform disposed between the fiber preform and the tool. The sacrificial preform is gas permeable. The process chamber is heated, and gaseous reactants are delivered into the process chamber during the heating. The gaseous reactants penetrate the through-holes of the tool and infiltrate the sacrificial preform and the fiber preform. Deposition of reaction products occurs on exposed surfaces of the fiber preform and the sacrificial preform, and a coating is formed thereon. In addition, the sacrificial preform accumulates excess coating material formed from increased reactions at short diffusion depths. Accordingly, the coating formed on the fiber preform exhibits a thickness variation of about 10% or less throughout a volume of the fiber preform.

A gas turbine engine includes a turbine section that includes a fan drive turbine. A geared architecture includes a sun gear in driving engagement with the fan drive turbine. A plurality of planet gears surrounds the sun gear. A ring gear surrounds the plurality of planet gears. A deflection limiter mechanically attaches the ring gear to an engine static structure. The deflection limiter includes a first support fixed to the ring gear that has a first interlocking feature and a second support fixed to the engine static structure that has a second interlocking feature. The first and second interlocking features define at least one of a radial clearance of between 0.005 inches (0.127 mm) and 0.080 inches (2.032 mm) or a circumferential clearance of between 0.005 inches (0.127 mm) and 0.250 inches (2.032 mm). A fan section includes a plurality of fan blades in driving engagement with the geared architecture through a fan drive shaft.

A gas turbine engine includes a turbine section that includes a fan drive turbine. A geared architecture includes a sun gear in driving engagement with the fan drive turbine. A plurality of planet gears surrounds the sun gear. A ring gear surrounds the plurality of planet gears. A deflection limiter mechanically attaches the ring gear to an engine static structure. The deflection limiter includes a first support fixed to the ring gear that has a first interlocking feature and a second support fixed to the engine static structure that has a second interlocking feature. The first and second interlocking features define at least one of a radial clearance of between 0.005 inches (0.127 mm) and 0.080 inches (2.032 mm) or a circumferential clearance of between 0.005 inches (0.127 mm) and 0.250 inches (2.032 mm). A fan section includes a plurality of fan blades in driving engagement with the geared architecture through a fan drive shaft.

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.

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.

A method of infiltrating a fiber preform comprises positioning an assembly in a process chamber, where the assembly includes a tool comprising through-holes, a fiber preform constrained within the tool, and a sacrificial preform disposed between the fiber preform and the tool. The sacrificial preform is gas permeable. The process chamber is heated, and gaseous reactants are delivered into the process chamber during the heating. The gaseous reactants penetrate the through-holes of the tool and infiltrate the sacrificial preform and the fiber preform. Deposition of reaction products occurs on exposed surfaces of the fiber preform and the sacrificial preform, and a coating is formed thereon. In addition, the sacrificial preform accumulates excess coating material formed from increased reactions at short diffusion depths. Accordingly, the coating formed on the fiber preform exhibits a thickness variation of about 10% or less throughout a volume of the fiber preform.

A cooled article and a method of forming a cooled article are disclosed. The cooled article includes a component, a porous material incorporated into the component, and a cooling medium within the porous material. Another cooled article is formed by a process includes the steps of forming a porous material from a pre-sintered preform, providing a component, and incorporating the porous material into the component. The porous material is in fluid communication with a cooling medium. The method of forming a cooled article includes providing a metal felt material infused with braze filler material, pre-sintering the metal felt material to form a porous material, providing a component, and incorporating the porous material into the component.

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.

A seal is formed of a matrix and includes hard particles. The matrix has a shear strength of greater than or equal to 200 psi and less than or equal to 2000 psi. A gas turbine engine is also disclosed.