Patent GB-789204-A (GB)
Title: Improvements in or relating to axial flow compressors or turbines
Inventor: BLYTH JACK VALLIS
IPC: F01D11/02
Filed: 1953-12-02
Abstract: 789,204. Turbines, gas-turbine plant axial-flow compressors. NAPIER & SON, Ltd., D. Oct. 27, 1954 [Dec. 2, 1953], No. 33538/53. Classes 110 (1) and 110 (3). A multi-stage axial flow compressor or turbine comprising at least two rotor discs each carrying a rotor blade ring and a stator blade ring between each adjacent pair of rotor blade rings has a diaphragm secured to the stator blade ring with sealing means at its inner diameter which engages a rotary sealing member mounted on the shaft, the diameter of the sealing member being considerably less than the diameter of the roots of the stator blades and means for supplying air under pressure to the annular spaces between the interstage body and the adjacent faces of the rotor discs through drillings or apertures in the hollow rotor shaft so as to provide a differential pressure acting on the rotor disc to oppose the axial thrust on the turbine rotor. The invention is described with reference to a propeller turbine engine in which an axial flow compressor supplies air to combustion equipment feeding a turbine driving the compressor; the forward end of which is connected through reduction gearing to a propeller. The turbine compressor shaft is hollow and is used to supply cooling air from immediately upstream of the last stage of the compressor to the turbine. Vanes on one of the compressor rotor discs act as a centripetal pump to assist the flow of cooling air. The turbine comprises three rotor discs 21, 22, 23, each carrying a ring of blades 25. Between the rings of blades 25 and upstream thereof there are arranged a number of rings of stator blades 26 with their outer ends anchored to the casing 27. The inner ends of the blades 26 are connected to interstage bodies 28, 29 formed of sheet metal. Each group of three adjacent stator blades have a common platform 30 which is connected to the body 28 or 29 by a floating tangential link 31 which allows for expansion of the blades and at the same time maintains the body 28 or 29 concentric with the turbine axis. The inner cylindrical surface of each interstage body has a labyrinth seal 32 which engages a cylindrical flange 33 on a sealing member arranged between each pair of adjacent rotor discs. Each sealing member has radial drillings or cutaway portions 34, 35 which communicate with radial holes 36 in the rotor shaft to permit cooling air to flow outwards from the shaft into the clearance spaces between each body 28, 29 and the adjacent rotor discs 21, 22, 23. The drillings 34 are larger than the drillings 35 so that the pressure on the downstream sides of the discs is higher than that on the upstream sides and thus part of the axial thrust on the turbine is balanced. The cooling air flow from the periphery of each clearance space is restricted by knife-edge seals 39. Each interstage body 28, 29 is formed of sheet metal discs 40, 41 welded to inner and outer annular rings 42, 43, and braced by radial spaces plates 44. The interior of the bodies 28, 29 is connected to a point midway along the labyrinth seals 32, 33 to reduce the pressure on the walls 40, 41. The external faces of the bodies 28, 29 and the stator wall 48 are provided with spiral fins 46 to cause the cooling air to circulate outwards in a spiral fashion over the face of the discs 21, 22, 23. A thrust balancing piston 50 is enclosed in a domed casing 52 to which air is admitted for the hollow turbine shaft. Detailed dimensions of the rotors, interstage bodies and spiral fins are given. The axial clearance between the interstage bodies and the rotor discs is preferably about one-eighth of the axial distance between the rotor discs. The diameter of the sealing member is preferably less than a half and may be one-third of the diameter of the blade root platforms associated with the interstage body. Specifications 751,010, 751,011 and 789,203 are referred to.