Nozzle Segment for a Gas Turbine Engine with Ribs Defining Radially Spaced Internal Cooling Channels

A gas turbine engine, comprising:

• a compressor:
• a combustor in flow communication with the compressor;
• and
• a turbine in flow communication with the combustor, the turbine including a turbine nozzle formed from an annular array of nozzle segments, each nozzle segment comprising:
• an airfoil formed from a ceramic matrix composite material, the airfoil including an exterior surface defining a pressure side and a suction side extending between a leading edge and a trailing edge, the airfoil defining an open internal volume within its interior for receiving a cooling medium, the open internal volume including a primary internal cavity and a plurality of internal cooling channels in flow communication with the primary internal cavity, the primary internal cavity extending within the interior of the airfoil from a location adjacent to the leading edge to a forward end of each of the plurality of internal cooling channels, the plurality of internal cooling channels extending within the interior of the airfoil from the primary internal cavity towards the trailing edge,
• wherein the plurality of internal cooling channels are spaced apart radially by a plurality of ribs extending within the interior of the airfoil.