1,091,621. Gas turbine and jet propulsion plant; centrifugal and axial-flow fans and compressors V. DAVIDOVIC. March 29, 1965, No. 13238/65. Headings F1C, F1G, F1J, F1L and F1T. [Also in Division F4] In a gas turbine engine, for stationary plant, automobiles, boats and aircraft, the turbine blades are cooled and the combustion air preheated, by passing some of the air discharged from the compressor 1 through a converging passage 13, passages 39, 41, the interior of hollow stationary blades 40 and rotary blades 43 of the turbine 3, passages 44, 45 and chamber 20 to the combustion chamber 2. Moreover, some of the air from the compressor passes through a branch pipe 16, scroll 18, passages 50, the interior of hollow stationary blades 51 and rotary blades 52 of the turbine 4, passages 54, 55 and chamber 20 to the combustion chamber. The air passing through the interior of the rotary blades is accelerated, due to centrifugal force, and at the exit of passages 45, 55 has an aspirating effect on the air flowing through adjacent passages 44, 54. The passages 44, 45, 54, 55 are formed by a number of rings 46, 56, 57 and longitudinal vanes 49 forming a grid, Fig. 5 (not shown). A portion of the air in scroll 18 passes through a number of rows of circumferentially-staggered tubes 95 forming a separate heat exchanger. Some of the air in passage 13 is delivered through a duct 25 and ports 24 into the combustion gases for cooling purposes. The combustion gases are exhausted through a manifold 26 and tail pipe 27. An internally finned chamber 14 receives liquid coolant, such as engine fuel, via connections 38 for cooling the air in passage 13. The inner wall of the passage 13 may be finned. Cooling of the air may be supplemented by injecting liquid, such as water, through circumferentially-spaced nozzles 58, the resulting vapour being carried by the air into the combustion chamber. The supply of liquid coolant is controlled by a carburetter (58'), Fig. 2 (not shown). Nozzles 59 alternating with nozzles 58, and nozzles 60 at the compressor may eject detergent into the flow passages. This ejection may be controlled manually or automatically at intervals. Gearing 7 may be provided for driving accessories 8 such as magnetoes, generators and pumps. In an alternative embodiment, Fig. 17 (not shown) some of the air passing through the branch pipe (136) is diverted through a combustion-gas-heated heat exchanger (112) (see Division F4) to the combustion chamber or chambers (104), and the air passing through the interior of the blades of both turbines is delivered to a downstream portion of the combustion chamber. Additionally the branch pipe is provided with a flap valve (139) for controlling the airflow therethrough, and is connected to a carburetter (141) for receiving liquid coolant. Nozzles (142, 143) supply either liquid coolant or detergent to the air flow upstream of both turbines. Cooling liquid may be supplied to cooling chambers (144, 145). In the embodiment in Fig. 3 (not shown), comprising a centrifugal compressor driven by one turbine and an axial-flow compressor driven by another turbine, an externally-finned connecting passage (13') between the two compressors, the axial-flow compressor and externally-finned ducts (81) between the latter compressor and the turbines, are surrounded by a jacket supplied with a liquid coolant by a pump mounted in the exhaust cone 84 and driven by gearing 85. In the embodiment in Figs. 8 to 11 (not shown), the external jacket (76) and the internal chamber (14') receive cooling air from a fan (88) upstream of the centrifugal compressor, the air being exhausted from the jacket and chamber via outlets (91, 92, 93) at the periphery of the engine casing. In a gas turbine jet propulsion engine for aircraft, Fig. 16 (not shown), an axialflow compressor delivers air to the combustion chamber or chambers (104) via a branch containing the hollow turbine blades, and a branch containing a combustion-gas-heated heat exchanger. Moreover, a fan upstream of the compressor discharges into the compressor, into a short duct (98) for additional propulsion, and into louvres (131) leading to a cooling jacket (130) surrounding the compressor. The air is withdrawn from the jacket by the suction produced by the air flow over the exterior of the engine casing. A turbine rotor, Figs. 13 to 15 (not shown), is constructed from a cast or forged hub (61) having internal vanes (62), cast, forged or bent channel members having blade parts (63, 64), and an outer shroud (68). Securing flanges (66) and abutting portions of the blade parts are welded together to form hollow blades. The shroud is formed by welding together two parts each having an outwardly extending rim (69). The rims (69) have either scolloped edges or tapered flutes (73) which, when the rotor is mounted in its casing, Fig. 7 (not shown), centrifugally impel any air leakage through openings (72) back into the main air flow. A similar sealing arrangement may be provided at the radially inner ends of the rotor blades. An annular deflector (74) adjacent the shroud deflects some of the combustion gases into the clearance about the downstream rim (69) to prevent air leakage. In an alternative sealing arrangement, Fig. 20 (not shown), the shroud is extended axially in both directions beyond the blades and is provided with labyrinth seals on its outer surface.