GB809514A - Improvements in or relating to combustion chambers - Google Patents

Abstract

809,514. Gas turbine plant. BRISTOL AERO-ENGINES Ltd. March 20, 1957 [April 3, 1956], No. 10215/56. Class 110 (3). [Also in Group XII] A combustion chamber 1 is of annular form, being disposed about an axial centre line 4 between a compressor 2 and turbine 3, the last two rows of the compressor rotor vanes 5, 6 and the last row of stator vanes 7 being shown. The combustion chamber comprises an inner wall 20 and an outer wall 21, an annular partition 22 being disposed between the inner and outer walls 20, 21 and extending from the inlet of the chamber to an intermediate part thereof. The partition 22 divides the upstream portion of the combustion chamber into an inner annular passage 26 through which dilution air flows from the compressor outlet through the inlet 24 with no reduction in its swirl velocity, and an outer annular space 25 which constitutes a combustion zone into which air flows from the compressor outlet through the inlet 23, likewise with no reduction in swirl velocity. The outer wall 21 is sharply divergent at 27 so that a toroidal eddy is formed at 28 into which fuel is injected through injector 29. The space 30 downstream of the partition constitutes a mixing zone in which dilution air from the passage 26 mixes with combustion products issuing from the combustion zone 25. The gaseous mixture flows between guide vanes 19 and impinges on turbine rotor blades 18. The outer wall 21 of the combustion chamber may be double-walled as shown, to permit flow of cooling air therethrough, the cooling air finally discharging through the hollow guide vanes 19 into an annular collector 31 and thence into the flow of combustion gases. The partition 22 is supported at its inlet end by means of vanes 33 which extend between the inner and outer walls 20, 21 and at its downstream end by means of vanes 34 which extend from the inner wall 20. The vanes 33, 34 and 19 do not greatly influence the swirl of gases passing therethrough so that the swirl of the air leaving the compressor remains substantially constant. The invention is also applicable in the case of an exhaust reheat combustion chamber of a gas turbine engine, in which case the exit of the chamber is wholly circular in form, and guide vanes at the chamber outlet are disposed so that the hot gases are discharged in an axial direction.