The present disclosure is directed to a gas turbine engine defining a radial direction, a longitudinal direction, and a circumferential direction, an upstream end and a downstream end along the longitudinal direction, and an axial centerline extended along the longitudinal direction. The gas turbine engine includes a fan assembly including a plurality of fan blades rotatably coupled to a fan rotor in which the fan blades define a maximum fan diameter and a fan pressure ratio. The gas turbine engine further includes a low pressure (LP) turbine defining a core flowpath therethrough generally along the longitudinal direction. The core flowpath defines a maximum outer flowpath diameter relative to the axial centerline. The gas turbine engine defines a fan to turbine diameter ratio of the maximum fan diameter to the maximum outer flowpath diameter. The fan to turbine diameter ratio over the fan pressure ratio is approximately 0.90 or greater.


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    Titel :

    Aircraft and direct drive engine under wing installation


    Beteiligte:

    Erscheinungsdatum :

    23.08.2022


    Medientyp :

    Patent


    Format :

    Elektronische Ressource


    Sprache :

    Englisch


    Klassifikation :

    IPC:    F02K JET-PROPULSION PLANTS , Strahltriebwerke / 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 / F02C Gasturbinenanlagen , GAS-TURBINE PLANTS



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