| 摘要 |
A blade 120 for a gas turbine engine comprises an aerofoil 122 having a root portion 124, a tip 126 portion located radially outwardly of the root portion, and leading and trailing edges 128, 130 extending between the root portion 124 and the tip portion 126. A shroud 132 extends transversely from the tip portion 126 of the aerofoil 122 and the aerofoil 122 defines interior cooling passages 140a-c which extend between the root portion 124 and the tip portion 126. The aerofoil 122 includes a wall member 142c adjacent the trailing edge 130 and a support structure 148 extending from the wall member 142c to the shroud 132 to support the shroud 132. The support structure 148 permits a flow of cooling air from a cooling passage 140a to the trailing edge 130 at a region proximate the tip portion 126 of the aerofoil 122. Optionally, the aerofoil 122 also includes a flow disrupting arrangement 160.
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