Abstract: A compressor of a gas turbine engine comprises blades 2, 4 provided on blade platforms 8, 12 of a rotor. Stator vanes 20 lie between the blades 2, 4 and are connected to a vane support structure 22. Sealing rings 38 are secured to the vane support structure 22 to restrict air flow into stator wells 36 on each side of the vane support structure 22. The sealing rings 38 are made from a flexible material so that they deflect on installation of the vane support structure 22 in a direction radially inwards between the adjacent blade platforms 8, 12.

Abstract: A compressor of a gas turbine engine comprises blades 2, 4 provided on blade platforms 8, 12 of a rotor. Stator vanes 20 lie between the blades 2, 4 and are connected to a vane support structure 22. Sealing rings 38 are secured to the vane support structure 22 to restrict air flow into stator wells 36 on each side of the vane support structure 22. The sealing rings 38 are made from a flexible material so that they deflect on installation of the vane support structure 22 in a direction radially inwards between the adjacent blade platforms 8, 12.

Abstract: A guide vane 1, such as may be used in a guide vane stage of a compressor (not shown), the guide vane comprising a main body 2 and a mounting base 4. In one aspect of the invention, at least part 22 of the mounting base 4 is aerofoil shaped. In another aspect of the invention, the mounting base 4 projects outwardly beyond a surface 14 of the main body 2 on one side only of the guide vane 1.

Abstract: A guide vane 1, such as may be used in a guide vane stage of a compressor (not shown), the guide vane comprising a main body 2 and a mounting base 4. In one aspect of the invention, at least part 22 of the mounting base 4 is aerofoil shaped.

Abstract: A gas turbine engine casing includes a cavity 10 separated from the main flow 12 through the engine by an annular array of tip treatment bars 16. The tip treatment bars 16 are hollow which enables the bars to deform or break upon impact by a shed blade 4 or blade fragment. This allows the blade or blade fragment to pass through the tip treatment bars 16. Blade or blade fragments may then be contained within the cavity 10, or embedded in the casing 2, so preventing further damage to the engine. Entire blades, or large blade fragments, may be able to break through the tip treatment bars 16 and through the engine casing 2 so as to leave the engine entirely, causing the minimum of damage. The interiors of the tip treatment bars may be filled with a damping material, to minimize high cycle fatigue failure.

Abstract: A variable camber vane for the inlet of a gas turbine engine has a pivoting trailing section behind a fixed leading section. In order to accommodate a wide range of flow conditions the trailing section is movable through a wide angle about its upstream edge. The surrounding casing profile and the radially outer edges of the vane trailing sections are matching part-spherical surfaces. As a result of the large angle of movement the trailing section is subjected to substantial turning forces. To contain the forces the curved radially outer edges of each trailing section is formed with a broad flange which tapers in the direction of the vane trailing edge, and for aerodynamic efficiency the flange is smoothly blended into the vane aerodynamic surfaces.