| 摘要 |
A blade for a gas turbine engine comprises an aerofoil having a root portion, a tip portion located radially outwardly of the root portion, and leading and trailing edges extending between the root portion and the tip portion. A shroud extends transversely from the tip portion of the aerofoil and the aerofoil defines interior cooling passages which extend between the root portion and the tip portion. The aerofoil includes a wall member adjacent the trailing edge and a support structure extending from the wall member to the shroud to support the shroud. The support structure permits a flow of cooling air from a cooling passage to the trailing edge at a region proximate the tip portion of the aerofoil. Optionally, the aerofoil also includes a flow disrupting arrangement.
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