GB814745A - An axial flow compressor - Google Patents

Abstract

814,745. Boundary layer control in axial flow compressors. UNITED AIRCRAFT CORPORATION. Jan. 5, 1956, No. 410/56. Class 4. [Also in Group XXVI] An axial flow compressor comprises a plurality of alternate rows of impeller and stator blades mounted respectively on a hub and a casing forming an annular passage having inner and outer walls, at least one row of blades having an annular scoop at the ends of the blades adjacent the hub for inducting the boundary layer flow on the inner wall, and radial passages within the blades of the row, communicating with the scoop and discharging adjacent the outer wall. In the part axial section of Fig. 2, each impeller disc 22 carries a row of blades 24 and an annular flange 40 extending upstream of the blades and forming a lip adapted to draw off the boundary layer from the inner shroud of the next upstream row of stator blades 34. The drawn off air is led to a chamber 44 and thence by a radial passage 46 to a downstream facing exhaust orifice 50 at the blade tip. The centrifugal action maintains the flow. In Fig. 4, lipped intakes 60 leading to chambers 64 are provided at the inner end of each row of stator blades the air discharging through passages 66 and 68 to the atmosphere at a comparatively low pressure. Similar lipped orifices 62 leading to chambers 70 and passages 72 are provided at the radially outer end of each row of stator blades. In Fig. 1 (not shown), the passages 68 discharge into the hollow double-walled casing of a gas turbine engine, and discharges thence from an annular nozzle surrounding the main turbine exhaust nozzle. This flow cools the combustion chambers in its passage from the slots 68 to the nozzle.