Patent GB-1363261-A (GB)
Title: Gas turbine engines
IPC: F04D27/02
Filed: 1971-08-17
Abstract: 1363261 Gas turbine ducted fan engines GENERAL ELECTRIC CO 17 Aug 1971 [20 Nov 1970] 38603/71 Headings F1C, F1G, F1J, F1L and FIT The invention relates to a gas turbine ducted fan engine and is described with reference to such an engine suitable for mounting in an aircraft to provide vertical lift. The combustion chamber is of toroidal form comprising a pair of liner walls 107, 108 which are spaced from the combustion chamber walls and which curve from a rearwardly facing fuel and air inlet forwardly, inwardly and rearwardly to a turbine nozzle guide vane assembly at the outlet of the combustion chamber. The engine shown in Fig. 3 comprises a contra-rotating axial-flow compressor 26, a toroidal combustion chamber 28, HP turbine 68 which drives the inner rotor 75 of the compressor and an LP turbine 40 which drives the outer rotary shell 54 of the compressor also the fan assembly 60. The HP turbine comprises a rotor 66 carrying blades 68 and formed with stub shafts 69, 70, the stub shaft 69 being connected to the rotor 75 of the compressor; five rows of blades 76 are mounted on the rotor. The HP turbine and compressor assembly is rotatably mounted on the shaft 44 of the LP turbine and compressor assembly by means of a forward bearing 92 and a rearward bearing 94. The LP turbine comprises three rotor discs 38 carrying blades 40 and connected to a rotor 42 which in turn is connected to the shaft 44. A compressor blade row assembly 45 is secured at the upstream end of the shaft, the assembly comprising blades 46, an inner shroud 48 and an outer shroud 50, the inner shroud being secured to the shaft 44. A cylindrical shell 54 is secured to the outer shroud by means of bolts 52, the shell comprising fan hub portion 55 and a plurality of stepped cylindrical portions. The hub portion 55 is formed with a series of axially-extending grooves in which the root portions of the fan blades 60 are mounted. Four rows of compressor blades 64 are mounted within the stepped portions of the shell by means of their shroud portions 81 which abut against each other and against the shell at the downstream end 91 and against the shroud 50 of the blade assembly 46 at the upstream end. The two turbine and compressor rotor assemblies rotate in opposite directions. Air entering the compressor at 26 is discharged through outlet guide vanes 97 to an annular diffuser 102. The air then passes in part around the liner wall 107 to the inlet 110 of the combustion zone and in part through passages in the turbine inlet guide vanes 109 to a space outwardly of the liner wall 108 and thence to the inlet 110. Combustion gases discharge through the guide vane assembly to act first on the HP turbine blades 68 and then on the LP turbine blades 40 finally discharging past the outlet guide blades 95 through a propulsion nozzle. The LP turbine and compressor assembly is rotatably mounted at its forward end in a bearing 78, the inner race of which is mounted on the sleeve 54, and at its rearward end in a bearing 80 which is supported in stator structure 90. The engine is started by supplying air under pressure through line 120 to a space 122 from which it is discharged through nozzles on to the HP turbine rotor blades 68. An oil reservoir is provided at 154 and in operation oil is pumped therefrom along the inner divergent surface of the shaft 44 and extension 156, the oil discharging through openings 158, 164, 166 in the shaft to lubricate the bearings 94, 92, also through openings 162 in the LP turbine rotor to lubricate the bearing 80. Oil also passes from the opening 166 through passages in the blades 46 and passages 172 in the hub 55, and openings  174 to lubricate the bearing 78; this oil returns to the reservoir through line 178. Air under pressure is tapped from an intermediate stage of the compressor through openings 124 in the shroud 81 and 126 in the shell 54, the air passing to a chamber 128 where it is utilized to prevent leakage through the seals 132, 134; air from the intermediate space between seals 132 passes through passages 136, 138 into the fan duct, and air from the intermediate space between seals 134 passes through duct 142 into the fan duct. In a second embodiment the compresor comprises two axial-flow stages and a centrifugal stage, the shell (254) being of reduced length.