A GDL cutting system of a fuel cell includes: a laser-cutting device that forms a gas diffusion layer by radiating a laser on the surface of a GDL fabric panel moving on a conveyer; an adsorbing-conveying device that adsorbs and conveys at least two gas diffusion layers cut by the laser-cutting device; a first vision sensor that senses an upper side of the gas diffusion layers cut by the laser-cutting device; and a second vision sensor that senses a lower side of the gas diffusion layers adsorbed and conveyed by the adsorbing-conveying device.

A method of machining at least one concrete part includes providing at least one concrete part; applying at least one marking to the at least one concrete part; arranging at least one machining device for machining the at least one concrete part relative to the at least one concrete part; detecting the at least one marking by at least one marking detection device and transmitting the marking data thus generated to the at least one machining device; and machining the at least one concrete part using the generated marking data by the at least one machining device.

A method of machining at least one concrete part includes providing at least one concrete part; applying at least one marking to the at least one concrete part; arranging at least one machining device for machining the at least one concrete part relative to the at least one concrete part; detecting the at least one marking by at least one marking detection device and transmitting the marking data thus generated to the at least one machining device; and machining the at least one concrete part using the generated marking data by the at least one machining device.

A gas turbine engine component includes a gas turbine engine component body formed of a ceramic matrix composite material having at least one fastener integrally formed with the gas turbine engine component body as a single-piece structure. The gas turbine engine component body initially comprises a rigidized preform structure formed from a polymer based material. The at least one fastener connects the gas turbine engine component body to an engine support structure.

A gas turbine engine component includes a gas turbine engine component body formed of a ceramic matrix composite material having at least one fastener integrally formed with the gas turbine engine component body as a single-piece structure. The gas turbine engine component body initially comprises a rigidized preform structure formed from a polymer based material. The at least one fastener connects the gas turbine engine component body to an engine support structure.

An airfoil for a gas turbine engine is provided that includes a first portion formed from a first plurality of plies of a ceramic matrix composite material and defining an inner surface of the airfoil, as well as a second portion formed from a second plurality of plies of a ceramic matrix composite material and defining an outer surface of the airfoil. The first portion and the second portion define a non-line of sight cooling aperture extending from the inner surface to the outer surface of the airfoil. In one embodiment, a surface angle that is less than 45° is defined between a second aperture and the outer surface. A method for forming an airfoil for a gas turbine engine also is provided.

An airfoil for a gas turbine engine is provided that includes a first portion formed from a first plurality of plies of a ceramic matrix composite material and defining an inner surface of the airfoil, as well as a second portion formed from a second plurality of plies of a ceramic matrix composite material and defining an outer surface of the airfoil. The first portion and the second portion define a non-line of sight cooling aperture extending from the inner surface to the outer surface of the airfoil. In one embodiment, a surface angle that is less than 45° is defined between a second aperture and the outer surface. A method for forming an airfoil for a gas turbine engine also is provided.

A method of manufacturing a honeycomb body, comprising extruding honeycomb extrudate (200) in an axial direction (A), the honeycomb extrudate (200) having an outer periphery (206); and laser machining in situ the honeycomb extrudate (200) to form a laser cut in the honeycomb extrudate. A system for in situ cutting a wet green ceramic extrudate, comprising a laser (500, 732, 826) configured to irradiate laser energy to an outer periphery of a wet green ceramic article, the laser energy adapted to cut through at least a portion of the outer periphery (206).

A method of manufacturing a honeycomb body, comprising extruding honeycomb extrudate (200) in an axial direction (A), the honeycomb extrudate (200) having an outer periphery (206); and laser machining in situ the honeycomb extrudate (200) to form a laser cut in the honeycomb extrudate. A system for in situ cutting a wet green ceramic extrudate, comprising a laser (500, 732, 826) configured to irradiate laser energy to an outer periphery of a wet green ceramic article, the laser energy adapted to cut through at least a portion of the outer periphery (206).

A method of forming a ceramic matrix composite component according to an exemplary embodiment of this disclosure, among other possible things includes laying up plies of ceramic reinforcement material with sacrificial plies to form a preform, infiltrating the preform with a ceramic matrix material, and machining away the sacrificial plies to reveal a surface profile of the ceramic matrix composite component. A preform for a ceramic matrix composite component is also disclosed.