TURBO-MACHINE STAGE Russian patent published in 2000 - IPC

FIELD: aircraft engine manufacture; compressors of gas-turbine engines. SUBSTANCE: turbo-machine stage has annular perforated passage wall with off-passage cavity and blading. Radial clearance is provided between blade ends and passage wall. Perforated passage wall has round holes arranged in two zones. Front zone is located in front of entrance edges of blades and hole axes of this zone are directed inwardly to gas flow towards front edges of blades at angle to generating line of passage wall in diametrical plane and in circular direction at angle equal to angle of flow at absolute motion in sectional area at blade entrance. Rear zone is located between entrance edge and blade-to-blade neck point on cascade back. Hole axes of rear zone are inwardly directed to off-passage cavity from gas-air passage at angle to generating line of passage wall in diametrical plane and in circular direction at absolute-flow angle in sectional area of blade-to-blade channel neck. Axial length of passage inner cavity equals at least total width of front and rear zones of holes and height of cavity is sufficient to provide annular flow area equal to at least that of all holes of rear zone. EFFECT: enlarged margin of gas-dynamics stability, reduced aerodynamic noise and vibration stresses in axial-flow compressor blades. 3 dwg