GB766812A - Improvements relating to multi-stage axial flow compressors - Google Patents

Abstract

766,812. Axial-flow compressors. ROLLSROYCE, Ltd. March 4, 1955 [March 4, 1954], No. 6413/54. Drawings to Specification. Class 110 (1). One or more rotor stages at the inlet end of a multi-stage axial-flow compressor include a blade or a group or groups of blades differing in aerodynamic characteristics or arranged to produce different aerodynamic effects as compared with the remaining blades of the stage, whereby the bands of high and low velocity air into which the air flow tends to break down when the stages are wholly or partially stalled are stabilized angularly with respect to the blade or group or groups of blades so as to rotate at the same speed as the rotor. The bands would otherwise have rotated at a reduced speed of, for example, 30 to 50 per cent of that of the rotor, thereby subjecting both the rotor and stator blading to exciting loads due to the succession of high and low air velocities which might have given rise to forced vibration of the blading. In one embodiment, each of the first two stages of the compressor of a gas-turbine jet-propulsion engine has two diametrically opposed groups of six consecutive rotor blades which have a greater camber than the remaining rotor blades of the respective stage. Additional stages, for example up to a total of five stages, may be so arranged. In all cases, it is arranged that air passing through blades of increased camber in one stage passes through like blades in succeeding stages. The blades may differ in aerodynamic characteristics other than camber, for example twist or angular setting. Groups of blades of the same or differing aerodynamic form as the standard blades may differ from the latter in respect of spacing.