GB753561A - Axial flow dynamic compressors, and gas turbine power plants utilising such compressors - Google Patents

Abstract

753,561. Axial-flow compressors; gas turbine and jet propulsion plant. PAVLECKA, V. H. Nov. 9, 1951 [May 25, 1951], No. 26364/51. Classes 110(1) and 110(3) An axial-flow compressor comprises a stationary contraprerotation stage, at least a first compression stage following the stationary stage and rotatable in the direction opposite to that of contrapreoration, at least a second compression stage rotatable in a direction opposite to that of the first compression stage, and means for rotating the compression stages. A gas turbine plant which may provide its output from the turbine shaft or, as shown in Fig. 2, may be used for jet propulsion, comprises a gas turbine 203 which directly drives one impeller 202 of the compressor and drives the other impeller 201 through planetary gearing 218 so that the impeller rotates in opposite directions at, preferably, equal speeds. Auxiliaries are driven through gearing 220, the shafts 44 of the auxiliaries passing through guide, blades 62. The air is accelerated by the guide blades to a subsonic or supersonic absolute velocity and enters the impeller 201 at a supersonic relative velocity, being subjected to an oblique, or oblique and reflected, or oblique and normal, or oblique, reflected and normal shocks. The air then enters the impeller 202 with a subsonic velocity and is subjected to further shocks, and enters the combustion chamber 200 with a substantially axial velocity. The combustion chamber comprises toroids 224, 226 of which the radial cross-sections are mutually eccentric so that the cross-sectional area of the flow channel 228 is substantially constant. A portion of the air flows through the channel 228 for cooling purposes, the remainder flowing through the toroid 226 by way of inlet and exit ports located in the radially outer part of the toroid and establishing a free vortex in the toroid. In a modification, orthodox combustion chambers are used, and the first impeller of the compressor is directly driven, the second impeller being oppositely rotated by planetary gearing located between the compressor and the turbine. Fig. 3 shows a compressor in which the rotor blades project into recesses 401, 402 which reduce leakage around the blade tips. The blading may be supersonic. In the plant shown in Fig. 4, the turbine has contrarotating stages 512, 514 and 526, 528, connected by planetary gearing 550. The contra-rotating impellers 506, 522 are driven from the turbine through cylindrical sleeves 510, 524. The toroidal combustion chamber 530 is carried by the sleeve 524 and consequently rotates. Fuel is supplied to nozzles 548 from a stationary pipe 540 by way of a rotating duct 542 and radial ducts 544. Instead of leading to a jet pipe, duct 553 may lead to a turbine or other consumer of compressed gas. In the plant shown in Fig. 9 compressor stages 1006, 1008 and 1010 and the combustion chamber 1022 are rotated by the turbine, while the remaining compressor stages are rotated in the opposite direction through planetary gearing 1004. A toroidal duct 1020 may be provided for the extraction of compressed air for external use. Fig. 7 shows a typical set of compressor blading, comprising inlet guide blades 801 followed by contra-rotating stages. Each blade 801 comprises a cylindrical surface 803, flat surfaces 804, 805, concave surface 806, and a convex surface 809, followed by supersonic nozzle surfaces 807, 808 and a flat surface 810. The air enters the guide blading at a subsonic velocity and leaves it at a supersonic velocity. The blades of the first two impellers are shaped for supersonic flow, with sharp leading edges, and comprise flat surfaces 811, 812 and a cylindrical surface 813. The first stage affords constant velocity flow for the air emerging from a compression shock 816, but the second stage may provide either constant velocity or diffusion. The blades of the last three stages are of normal form, for subsonic flow. In the blading shown in Fig. 10, the inlet guide blades have cylindrical leading edges 1102, flat surfaces 1103, 1104 and cylindrical surfaces 1105, 1109. The air is accelerated by the inlet guide blades to a velocity which is subsonic in absolute value and supersonic with respect to the blades of the first impeller. An oblique shock 1110 occurs as in Fig. 7. The air enters the second impeller at a supersonic velocity ana, after an oblique shock 1117, passes at subsonic velocity through a diffusing passage formed by diverging surfaces 1118, 1120. The inlet guide blades may be of the supersonic form shown in Fig. 7, and there may be more than one supersonic impeller. The inlet guide blades 1901 shown in Fig. 16 are only slightly curved, and the air flows between them with substantially constant velocity. As before, the relative velocity at entry to the blades of the first and second impellers is' supersonic, and oblique shocks 1902, 1907 occur. Diffusion occurs partly in the subsonic portion of impeller 1906 but mainly in a stationary diffuser 1908. Specifications 743,474, 743,475 [Group III] and 753, 652 are referred to.