GB887083A - Improvements relating to axial flow compressors - Google Patents

Abstract

887,083. Axial-flow compressors and turbines. ROLLS-ROYCE Ltd. July 15, 1960 [Aug. 5, 1959], No. 26726/59. Drawings to Specification. Classes 110 (1) and 110 (3). In a multi-stage axial-flow compressor for a gas turbine engine, the final rotor stage is designed for variable work with blade height, whereby the compressor total pressure distribution is matched to the total pressure distribution required by a combustion chamber to provide good combustion in the gas turbine engine. For example, the outlet total pressure may be made constant over the blade height or be given a minimum value at blade mid-height with maximum values near root and tip. Assuming the static pressure to be substantially constant over the blade height, departure of the total pressure from constancy is due to reductions of the axial component of velocity caused by losses at root and tip and can be corrected by increasing the rotor blade exit angle at blade midheight and, it may be, decreasing this angle at root and tip so as to produce a more constant distribution of the axial component of velocity. The variation of the rotor blade exit angle also causes the absolute velocity at entry to the outlet stator blades to be greater at root and tip than at blade mid-height; this allows for the losses at root and tip and permits the axial exit velocity from the outlet stator to be more nearly constant radially. According to the Provisional Specification, the invention may be applied to compressors for other uses, for example, for supplying air to helicopter rotors of the type driven by air jets, or to closed circuit wind tunnels. The invention is also applicable to turbines, particularly in gas turbine engines using reheat.