Patent ID: 8740551

Claim:
A blade outer air seal assembly of a gas turbine engine having a main axis of rotation defining axial, radial and circumferential directions, the blade outer air seal assembly comprising: an array of circumferentially adjoining blade outer air seal support segments forming a support ring; and an array of circumferentially adjoining blade outer air seal segments forming a static turbine shroud surrounding a turbine rotor, the turbine shroud being supported within the support ring, the blade outer air seal segments each including a platform extending axially from a leading edge to a trailing edge and circumferentially between opposed circumferential sides of the platform, front and rear hooks to support the platform radially and inwardly spaced apart from the support ring, to thereby define an annular cavity between the front and rear hooks, the platform defining a plurality of cooling passages extending axially through the platform, each of the cooling passages having an inlet defined in a radially outer surface of the platform and an exit defined on the leading edge, each inlet communicating with the annular cavity for intake of cooling air from the annular cavity, and wherein at least one of said cooling passages positioned close to the respective opposed circumferential sides extends linearly from one of the inlets and is circumferentially skewed away from the axial direction such that the at least one of said cooling passages has said inlet circumferentially spaced apart from an axial seal slot defined in the opposed circumferential sides of the platform and has an exit hole circumferentially aligned with said axial seal slot to thereby cool a corner area of the platform between the leading edge and the respective opposed circumferential sides.