GB936635A - Multi-stage axial-flow compressor - Google Patents

Abstract

936,635. Axial-flow compressors; compressor blade mountings. ROLLS-ROYCE Ltd. April 13, 1962 [April 21, 1961], No. 14616/61. Classes 110(1) and 110(3). A multi-stage axial-flow compressor has a bleed air passage through which some of the air compressed by the compressor may be bled from one of the compressor stages, which stages have stator blades each of whose platforms is provided with at least one surface which is so inclined with respect to the compressor axis as to direct some of the compressed air radially outwardly and in a downstream direction towards the bleed air passage. A gas turbine engine (not shown) for use on an aircraft (not shown) comprises a multi-stage compressor 10 having a casing 11 which includes annular U- section wall members 12, 13, 14. The platforms 32 of the stator blades of the sixth compressor stage have internal surfaces 33, which are inclined at an angle of 40 to 50 degrees to the compressor axis, to direct some of the compressed air radially outwardly and in a downstream direction so as to pass through angularly spaced elongated slots 25, 24 into a bleed air passage 36. This air may, for example, be used for the operation of blown flaps on the aircraft. Tapered ribs 37, Fig. 3, formed between the slots 25 are rounded at their upstream ends to reduce pressure losses. Means (not shown) may be provided for closing the slots 25, 24 or preventing flow through the bleed air passage 36. In an alternative construction (not shown), the compressor casing has an integral radially outwardly extending projection which comprises a short axially extending wall connected with inclined upstream and downstream walls. The platforms of the stator blades, are secured parallel to the upstream wall which is radially outwardly inclined in a downstream direction; the inclined downstream wall is provided with angularly spaced elongated slots. Separate inner wall members support the stator blades of adjacent stages.