Patent GB-1099677-A (GB)
Title: Improvements in axial-flow turbines and compressors
Inventor: RICHARDSON PAUL ERNEST COLIN
IPC: F01D17/16
Filed: 1963-11-29
Abstract: 1,099,677. Axial-flow compressors and turbines; jet propulsion engines. BRISTOL SIDDELEY ENGINES Ltd. Nov. 19, 1964 [Nov. 29, 1963], No.47274/63. Addition to 992,941. Headings F1C, F1J and FIT. In an axial-flow turbine or compressor for use with outlet ducting defined at least partly by a surface which is non-coaxial with respect to the rotor axis so as to be capable of causing a non- uniform distribution of the flow of working fluid through the rotor blades liable to cause damaging vibration of the rotor blades, at least a portion of at least one blade in a final ring of stator blades is pivotally mounted to permit individual adjustment of its exit angle, to enable the vibration to be reduced or eliminated. Fig. 4 shows part of a gas turbine engine, in which the exhaust duct 31 is branched to provide two lateral outlets 34 each having a rotatable gas deflecting nozzle 35. A bearing housing 33 is supported by a small number, say seven, of large vanes 32. To compensate for the variation of swirl in the exhaust gases leaving the last ring of turbine rotor blades 28 caused by variation in operation conditions of the engine, a row of variable-pitch stator blades 37 can be turned in unison in accordance with an operating condition of the engine, such as rotor speed, by means of a control ring 43 and arms 41. The stator blades 37 reduce the swirl angle to a value which can be dealt with satisfactorily by the relatively few vanes 32. The trunnion 40 of each stator blade 37 is adjustably attached to its arm 41 so that the blades 37 can be initially and individually adjusted in pitch whereby the pitch varies from blade to blade in such manner as to reduce or eliminate the tendency for the rotor blades 28 to be set into vibration by the non- uniformity of circumferential distribution of flow caused by the vanes 32, the branched shape of the exhaust duct 31 or the presence of the nozzles 35. Fig. 3 shows part of a by-pass type gas turbine engine in which individually-adjustable (but otherwise fixed) stator blades 19 are provided to compensate for the non-uniformity of circumferential distribution of flow in the by-pass annulus 16 tending to set the final rotor blades 10 of the lowpressure compressor into vibration. The non- uniformity may be caused by vanes 9a, which are few in number, say four, or by non-coaxial outlets. A square-section or splined extension 21 of the stem 20 of each blade 19 carries a vernier plate 22 seating on the casing 1, and the blade is locked by a nut 23 after adjustment. The blades 19, which preferably have no camber or twist, may be more or less equally distributed around the annulus 16 or may be provided in one or more groups each in a region where flow correction may be required. For example, they may be arranged in two arcuate groups at the top and bottom of the annulus 16 to compensate for the effect of an outlet opening at each side of the engine. Non-adjustable or adjustable stator blades 18 may be arranged with a circumferential variation of turning effect on the flow to offset the effect of flow disturbances caused by vanes 9b and, it may be, vanes 9a also. Adjustable flow-compensating stator blades provided at the rear of the compressor of a gas turbine engine may each have a non-adjustable leading edge part and an adjustable trailing edge part.