Abstract: A nozzle guide vane structure for a gas turbine engine in which the vane comprises an aerofoil portion and inner and outer platforms. The non gas-contacting surfaces of the platforms are provided with forward and rearward sealing flanges sealed to adjacent engine structure to provide spaces sealed against ingress of working fluid between each pair of forward and rearward flanges. The mounting means for the vane projects into these spaces from the non gas-contacting surface.

Abstract: To minimize leakage of high pressure compressor air from the high pressure side of an aixal flow compressor rotor back to the low pressure side by the outside of a ring of shroud segments, the shroud segments, or at least their front and rear edge portions, are made relatively thin and springy and are retained to a carrier ring so that those edge portions are held sprung against circumferentially extending sealing lands on the supporting structure. It is arranged that the bending in the shroud segments is effective to accommodate predetermined manufacturing tolerances in the shroud segments and the supporting structure, thus avoiding unwanted clearances at the sealing lands. Alternatively, or in addition, the bending can accommodate reductions in bending due to differential expansion between the components during operation of the compressor rotor at elevated temperatures.

Abstract: An aerofoil member such as a blade or vane for a gas turbine engine comprises a hollow aerofoil with a plug of porous material filling a region of the hollow interior. The porous material forms at least part of the external surface of the aerofoil so that cooling fluid can flow from the interior of the aerofoil through the plug to the external surface of the aerofoil. The plug is preferably at the leading or trailing edge of the aerofoil. Methods of making the member include investment casting using a porous ceramic core to form the porous plug, or using sintered particulate material to form the porous plug.

Abstract: A blade or vane structure for a gas turbine engine comprises a hollow aerofoil member and a cooling air entry tube within the hollow aerofoil member, the aerofoil and the tube comprising a unitary cast structure. To make this structure it is proposed that a lost wax casting process should be utilized, with the internal cavities formed using ceramic cores formed with the aid of temporary, disposable pieces which define the space eventually occupied by the tube.

Abstract: A blade or vane structure for a gas turbine engine comprises a hollow aerofoil member and a cooling air entry tube within the hollow aerofoil member, the aerofoil member and the tube comprising a unitary cast structure. To make this structure it is proposed that a lost wax casting process should be utilized, with the internal cavities formed using ceramic cores formed with the aid of temporary, disposable pieces which define the space eventually occupied by the tube.

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.

Abstract: A jet pipe wall is made from a honeycomb sandwich structure in which the skin forming one side of the sandwich is cut into many small square sheet pieces so that when differential longitudinal expansion occurs between the skins as a result of lying in fluid flows of differing temperature, the total shear felt by the honeycomb is divided equally between those areas of honeycomb covered by respective small sheet pieces, so that the overall shear effect is reduced.