Abstract: In order to facilitate efficient use of high pressure cooling air in a turbine rotor blade its aerofoil has a triple-pass convoluted cooling air duct in its leading edge region, and a triple-pass convoluted cooling air duct in its mid-chord region, both of them being fed from a common high pressure inlet in the root of the blade. The trailing edge region has a single-pass duct fed by low pressure cooling air from an inlet located just under the blade platform. The shroud of the blade also has cooling air passages and these are fed from the ends of respective ones of the ducts. Air from the ducts also exits through rows of film cooling holes to film cool the concave flank of the aerofoil and through a row of holes in its trailing edge to remove heat from the thinner metal section in that area.

Abstract: The combustion equipment of a gas turbine engine comprises an annular flame tube and a support casing both subjected to relatively high temperatures. Engine acceleration or engine deceleration causes the temperature in the combustion equipment to increase or decrease respectively, and this results in either expansion or contraction of the annular flame tube and the support casing.The annular flame tube has a greater rate of thermal expansion/contraction than that of the support casing and this results in distortion or cracking of the combustion equipment if the annular flame tube and support casing are rigidly mounted to each other.

Abstract: A nozzle guide vane assembly for a gas turbine engine comprises a circumferentially extending array of angularly spaced apart aerofoils each having projections adapted to engage with structure to retain the aerofoil in its longitudinal direction. Inner and outer platform members are separate from the aerofoils and each comprises a thicker support skin and a thinner inner skin. Both skins have aerofoil shaped apertures through which the aerofoils project, the support skin retaining the aerofoil against twisting, circumferential or axial loads and the inner skin serving to define a respective boundary of the gas flow through the assembly. The aerofoil is free to slide through the apertures sufficiently to permit relative expansion in a direction longitudinal of the aerofoil, and sealing means are associated with the inner skins and provide a seal between the skins and the aerofoils.