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.


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    Titel :

    Gas turbine engine


    Beteiligte:
    CLARE BEWICK (Autor:in) / STEVEN RADOMSKI (Autor:in)

    Erscheinungsdatum :

    2021-05-26


    Medientyp :

    Patent


    Format :

    Elektronische Ressource


    Sprache :

    Englisch


    Klassifikation :

    IPC:    F02C Gasturbinenanlagen , GAS-TURBINE PLANTS / B64D Ausrüstung für Flugzeuge , EQUIPMENT FOR FITTING IN OR TO AIRCRAFT / F01D Strömungsmaschinen [Kraft- und Arbeitsmaschinen oder Kraftmaschinen], z.B. Dampfturbinen , NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES



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