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A geared turbofan engine includes a first rotor, a fan, a second rotor, a gear train, a fan casing, a nacelle and a plurality of discrete acoustic liner segments. The fan is connected to the first rotor and is capable of rotation at frequencies between 200 and 6000 Hz and has a fan pressure ratio of

A geared turbofan engine includes a first rotor, a fan, a second rotor, a gear train, a fan casing, a nacelle and a plurality of discrete acoustic liner segments. The fan is connected to the first rotor and is capable of rotation at frequencies between 200 and 6000 Hz and has a fan pressure ratio of between 1.25 and 1.60. The gear train connects the first rotor to the second rotor. The fan casing and nacelle are arranged circumferentially about a centerline and define a bypass flow duct in which the fan is disposed. The plurality of discrete acoustic liner segments with varied geometric properties are disposed along the bypass flow duct.

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1. A geared turbofan engine, comprising: a first rotor;a fan connected to the first rotor, wherein the fan is capable of rotation at frequencies between 200 and 6000 Hz and has a fan pressure ratio of between 1.25 and 1.60;a second rotor;a gear train that connects the first rotor to the second rotor

1. A geared turbofan engine, comprising: a first rotor;a fan connected to the first rotor, wherein the fan is capable of rotation at frequencies between 200 and 6000 Hz and has a fan pressure ratio of between 1.25 and 1.60;a second rotor;a gear train that connects the first rotor to the second rotor;a fan casing and a nacelle arranged circumferentially about a centerline and defining a bypass flow duct in which the fan is disposed; anda plurality of discrete acoustic liner segments having varied geometric properties disposed along the bypass flow duct; wherein the plurality of discrete acoustic liner segments comprises a first acoustic liner segment, and a second acoustic liner segment spaced apart from the first acoustic liner segment;wherein each the first and second acoustic liner segments includes a cellular core structure, the cellular core structure comprising one or more resonator chambers having a width; andwherein the width of the one or more resonator chambers of the first acoustic liners differs from the width of the one or more resonator chambers of the second acoustic liner. 2. The geared turbofan engine of claim 1, wherein at least one discrete acoustic liner segment of the plurality of discrete acoustic liner segments is disposed on an inner surface of the nacelle inside the bypass flow duct. 3. The geared turbofan engine of claim 1, wherein at least one discrete acoustic liner segment of the plurality of discrete acoustic liner segments is disposed on an inner surface of the fan casing inside the bypass flow duct. 4. The geared turbofan engine of claim 1, wherein the gas turbine engine further comprises: a core casing arranged circumferentially around the centerline within the nacelle and the fan casing and defining an inner surface of the bypass flow duct; andwherein at least one discrete acoustic liner segment of the plurality of discrete acoustic liner segments is disposed on the inner surface of the bypass flow duct. 5. The geared turbofan engine of claim 1, wherein the cellular core structure of one of the plurality of discrete acoustic liners has a depth that differs from a depth of the cellular core structure of another of the plurality of discrete acoustic liner segments. 6. The geared turbofan engine of claim 1, wherein a cross-sectional geometry of the one or more resonator chambers of one of the plurality of discrete acoustic liner segments differs from a cross-sectional geometry of another of the one or more resonator chambers. 7. The geared turbofan engine of claim 1, wherein each of the plurality of discrete acoustic liner segments includes a face sheet with holes and the holes communicate with the resonator chambers in the cellular core structure, wherein a diameter of the holes in the face sheet of one of the plurality of discrete acoustic liner segments differs from a diameter of holes in the face sheet of another of the plurality discrete acoustic liner segments. 8. The geared turbofan engine of claim 7, wherein a number of the holes in the face sheet of one of the plurality of discrete acoustic liner segments differs from a number of the holes in the face sheet of another of the plurality of discrete acoustic liner segments. 9. The geared turbofan engine of claim 7, wherein a thickness of the face sheet of one of the plurality of discrete acoustic liner segments differs from a thickness of the face sheet of another of the plurality of discrete acoustic liner segments. 10. The geared turbofan engine of claim 7, wherein the face sheet of at least one of the discrete acoustic liner segments is micro-perforated. 11. A geared turbofan engine, comprising: a gear train connecting a first rotor to a second rotor;a core casing extending circumferentially around the first rotor and defining a portion of an inner surface of a bypass flow duct;a nacelle and a fan casing extending circumferentially around the core casing and defining an outer surface of the bypass flow duct; andan acoustic liner with two or more zones disposed along the bypass flow duct, the two or more zones being tuned to attenuate a different frequency range of acoustic noise;wherein a first zone of the two or more zones comprises a first face sheet having a first radial thickness; andwherein a second zone of the two or more zones comprises a second face sheet having a second radial thickness different from the first radial thickness. 12. The geared turbofan engine of claim 11, wherein the gear train comprises an epicyclic transmission. 13. The geared turbofan engine of claim 11, wherein the geared turbofan further comprises: a fan connected to the first rotor; anda low speed spool driving the second rotor, the low speed spool including a low pressure compressor section and a low pressure turbine section. 14. The geared turbofan engine of claim 13, wherein the fan rotates at frequencies under 1000 Hz and one of the zones of the acoustic liner is tuned to attenuate frequencies under 1000 Hz. 15. The geared turbofan engine of claim 13, wherein one of the zones of the acoustic liner is tuned to attenuate frequencies above 1000 Hz. 16. The geared turbofan engine of claim 11, wherein the acoustic liner is segmented into discrete axial segments. 17. The geared turbofan engine of claim 11, wherein the acoustic liner is segmented into discrete circumferential segments. 18. The geared turbofan engine of claim 11, wherein the acoustic liner is segmented into discrete segments and each discrete segment contains a single zone of the multiple zones. 19. The geared turbofan engine of claim 11, wherein the acoustic liner is segmented into discrete segments and at least one discrete segment contains more than one zone of the multiple zones. 20. A gas turbine engine, comprising: a fan rotatably arranged along an axial centerline;a fan casing and a nacelle arranged circumferentially around the centerline and defining a bypass flow duct in which the fan is disposed; anda plurality of discrete acoustic liner segments with varied geometric properties disposed along the bypass flow duct; wherein the plurality of discrete acoustic liner segments comprises a first acoustic liner segment, and a second acoustic liner segment spaced apart from the first acoustic liner segment;wherein each the first and second acoustic liner segments includes a cellular core structure, the cellular core structure comprising one or more resonator chambers having a width; andwherein the width of the one or more resonator chambers of the first acoustic liners differs from the width of the one or more resonator chambers of the second acoustic liner. 21. The gas turbine engine of claim 20, wherein at least one discrete acoustic liner segment of the plurality of discrete acoustic liner segments is disposed on an inner surface of the nacelle inside the bypass flow duct. 22. The gas turbine engine of claim 20, wherein at least one discrete acoustic liner segment of the plurality of discrete acoustic liner segments is disposed on an inner surface of the fan casing inside the bypass flow duct. 23. The gas turbine engine of claim 20, wherein the gas turbine engine further comprises: a core casing arranged circumferentially around the centerline within the nacelle and the fan casing and defining an inner surface of the bypass flow duct; andwherein at least one discrete acoustic liner segment of the plurality of discrete acoustic liner segments is disposed on the inner surface of the bypass flow duct. 24. The gas turbine engine of claim 20, wherein the cellular core structure of one of the plurality of discrete acoustic liners has a depth that differs from a depth of the cellular core structure of another of the plurality of discrete acoustic liner segments. 25. The gas turbine engine of claim 24, wherein a cross-sectional geometry of the one or more resonator chambers of one of the plurality of discrete acoustic liner segments differs from a cross-sectional geometry of another of the one or more resonator chambers. 26. The gas turbine engine of claim 24, wherein each of the plurality of discrete acoustic liner segments includes a face sheet with holes and the holes communicate with the resonator chambers in the cellular core structure, wherein a diameter of the holes in the face sheet of one of the plurality of discrete acoustic liner segments differs from a diameter of holes in the face sheet of another of the plurality discrete acoustic liner segments. 27. The gas turbine engine of claim 26, wherein a number of the holes in the face sheet of one of the plurality of discrete acoustic liner segments differs from a number of the holes in the face sheet of another of the plurality of discrete acoustic liner segments. 28. The gas turbine engine of claim 26, wherein a thickness of the face sheet of one of the plurality of discrete acoustic liner segments differs from a thickness of the face sheet of another of the plurality of discrete acoustic liner segments. 29. The gas turbine engine of claim 26, wherein the face sheet of at least one of the discrete acoustic liner segments is micro-perforated. 30. A gas turbine engine, comprising: a core casing extending circumferentially around the first rotor and defining a portion of an inner surface of a bypass flow duct;a nacelle and a fan casing extending circumferentially around the core casing and defining an outer surface of the bypass flow duct;a fan rotatably disposed in the bypass flow duct; andan acoustic liner with two or more zones disposed in the bypass flow duct, wherein the two or more zones being tuned to attenuate a different frequency range of acoustic noise;wherein a first zone of the two or more zones comprises a first face sheet having a first radial thickness; andwherein a second zone of the two or more zones comprises a second face sheet having a second radial thickness different from the first radial thickness. 31. The gas turbine engine of claim 30, wherein the fan rotates at frequencies under 1000 Hz and one of the zones of the acoustic liner is tuned to attenuate frequencies under 1000 Hz. 32. The gas turbine engine of claim 31, wherein one of the zones of the acoustic liner is tuned to attenuate frequencies above 1000 Hz. 33. The gas turbine engine of claim 30, wherein the acoustic liner is segmented into discrete axial segments. 34. The gas turbine engine of claim 30, wherein the acoustic liner is segmented into discrete circumferential segments. 35. The gas turbine engine of claim 30, wherein the acoustic liner is segmented into discrete segments and each discrete segment contains a single zone of the multiple zones. 36. The gas turbine engine of claim 30, wherein the acoustic liner is segmented into discrete segments and at least discrete segment contains more than one zone of the multiple zones.