GB768370A - Improvements in gas turbine engines, particularly for aircraft propulsion - Google Patents

Abstract

768,370. Gas turbine engines. SOC. NATIONALE D'ETUDE ET DE CONSTRUCTION DE MOTEURS D'AVIATION. July 27, 1954 [Aug. 18, 1953], No. 21853/54. Class 110 (3). [Also in Group XII] In a gas turbine engine, at least one small auxiliary combustion chamber is disposed within a combustion zone of the main combustion equipment, the auxiliary chamber being supplied with air at a pressure greater than the pressure in the main combustion zone by means of an auxiliary compressor which takes air tapped from the main compressor of the engine and compresses it to a higher pressure. The auxiliary chamber has discharge passages into the main combustion zone. The gas turbine engine shown in Fig. 1 comprises an air compressor 1 driven by a turbine 4 which is supplied with gases from combustion equipment 5. Air tapped from the compressor discharge passes through a duct 14 and is further compressed in a centrifugal compressor 13. This further compressed air passes through ducts 16 into a series of auxiliary combustion chambers 9 disposed within the main combustion equipment 5 which comprises an annular flame tube. Fuel is supplied to the auxiliary combustion chambers 9 through injectors 12 and the gases issue through outlets 19 into the main combustion zone. The auxiliary chambers 9 are disposed between main fuel injectors 10 which are provided with baffles 11. The compressor 13 is driven by an auxiliary gas turbine 20 supplied with gases tapped from the main combustion equipment 5 through duct 21. The compressor 13 may alternatively be driven by gearing from the main turbine and in one embodiment it is rotatably mounted on the turbine-compressor shaft and is driven therefrom by epicyclic gearing. In a further embodiment, the main combustion equipment comprises separate combustion chambers and an auxiliary combustion chamber 9 is disposed within each combustion chamber.