A gas turbine engine 10 for an aircraft, comprising an engine core 11 surrounded by an inner nacelle 53, a fan 23 housed in a fancase 52. A bypass duct (22, fig1) is defined between the inner nacelle and an outer nacelle 21. The fancase is connected to the upstream end of the engine core by a circumferential array of radial guide vanes 25. The outer nacelle is connected to a downstream end of the engine core by a single bifurcation 29. Having a single bifurcation could reduce the number of obstructions in the bypass duct and could improve the specific fuel consumption of the engine. A drainage path 33 could be used for draining flammable fluid from the engine core through a drainage guide vane 51 of the circumferential array of guide vanes 25. A second drainage path 39 could be used, which slopes downwardly from an upstream end of the engine core to a discharge outlet 53.
Gas turbine engine
2021-05-26
Patent
Electronic Resource
English