Patent GB-980306-A (GB)
Title: Gas turbine engine
Inventor: HEWSON COLIN TAYLOR
IPC: F02K3/075
Filed: 1963-05-10
Abstract: 980,306. Gas turbine jet propulsion plant; axial-flow compressors and fans. ROLLSROYCE Ltd. Sept. 17, 1963 [May 10, 1963], No. 36650/63. Divided out of 978,658. Headings F1C and F1J. [Also in Division B7] In a gas turbine engine having an axial-flow compressor, combustion equipment, a turbine and an exhaust duct, all arranged in flow series, the compressor comprises at least one row of rotor blades driven by the turbine and at least one row of rotatable stator blades mounted adjacent to the, or a, row of rotor blades, the construction being such that the rotatable stator blades may, in operation, be rotated at a tip speed which is less than, but is related to, that of the rotor blades, and the engine comprises a fan having a duct of which the upstream end is arranged to receive ram air and having at least one row of fan rotor blades arranged in the fan duct and driven by the row or rows of rotatable stator blades, the pressure ratio across the fan being substantially the same as that across the compressor. The gas turbine aircraft engine shown in Fig. 1 comprises an L.P. compressor 11 and an H.P. compressor 12, driven through concentric shafts 23, 21 by an L.P. turbine 15 and an H.P. turbine 14, respectively. The stator blade rows 24 of the L.P. compressor are carried by a rotatably mounted annular casing 25. The air passing through the L.P. compressor rotates the stator blades at a tip speed which is less than, but is related to, that of the rows 22 of rotor blades, whereby the relative tip speed of each row of rotor blades with respect to the adjacent row of rotatable stator blades is reduced to a subsonic value. The rotatable annular casing 25 also carries rows 31 of rotor blades of a fan 30, which are therefore driven by the rows 24 of rotatable stator blades. The fan duct 26 dissharges into a by-pass duct 17 which also receives part of the air compressed by the L.P. compressor 11. The by-pass duct 17 discharges through a number of angularly spaced mixer shoots into the engine exhaust duct 16. A row of pivotally mounted inlet guide vanes 32 and a valve mechanism 33 may be provided, coupled to a common control shaft so that opening movement of the guide vanes 32 is associated with closing movement of the valve mechanism 33 and vice versa. The control may be operated to reduce the flow through the fan 30 at starting, whereby the fan speeds up, thereby reducing the relative speed between the rows 22 and 24 of L.P. compressor blades. Consequently, the L.P. compressor 11 does less work, whereby the flow demanded by the H.P. compressor 12 will not be too small for the L.P. compressor, whereby surging of the latter at starting is avoided. The valve mechanism 33 and the mixer shoots 20 may be omitted, the by-pass duct then extending to the downstream end of the exhaust duct without communicating with it. In this case an I.P. compressor and an I.P. turbine may be provided, both mounted on the L.P. shaft. The separate turbines may be replaced by a single-shaft turbine driving both compressors, and the provision of variable inlet guide vanes 32 for facilitating starting may then be unnecessary. Fig. 5 shows an engine in which the L.P. compressor 11 e does not communicate with the by-pass duct 17e. To facilitate starting variable inlet guide vanes 32e are closed and a brake 37e is applied to prevent rotation of rows 24e or rotatable stator blades and rows 31e  of fan rotor blades. At take-off, the inlet guide vanes 32e are adjusted to give a high by-pass ratio; thereafter, they may be adjusted to reduce the by-pass ratio to a relatively low value. At high speed (e.g., Mach 3 or Mach 4), the brake 37e is again applied to cause the engine to operate as a non-by-pass engine.