Acoustic liner with varied properties

Abstract

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.

The invention claimed is: 
     
       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; and   a 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; and   wherein 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; and   wherein 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; and   an 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; and   wherein 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; and   a 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; and   a 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; and   wherein 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; and   wherein 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; and   an 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; and   wherein 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. 
     
     
       37. A geared turbofan engine comprising:
 an engine core comprising:
 a first rotor connected to a fan; 
 a second rotor; and 
 a gear train connecting the first rotor to the second rotor; 
   a core casing disposed circumferentially around at least a portion of the engine core;   a nacelle disposed circumferentially around at least a portion of the core casing, wherein a bypass flow duct is defined between the nacelle and the core casing; and   an acoustic liner extending at least partially around a circumference of the bypass flow duct and disposed on an inner surface of the nacelle, the acoustic liner comprising:
 a cellular core; and 
 a face sheet disposed on the cellular core and defining a surface of the bypass flow duct; 
   wherein a first circumferential zone of the acoustic liner extends around a first portion of the circumference of the bypass flow duct, in which the cellular core of the acoustic liner in the first circumferential zone includes multiple, circumferentially adjacent resonator chambers each having a first depth; and   wherein a second circumferential zone of the acoustic liner extends around a second portion of the circumference of the bypass flow duct, in which the cellular core of the acoustic liner in the second circumferential zone includes multiple, circumferentially adjacent resonator chambers each having a second depth different from the first depth.   
     
     
       38. The geared turbofan engine of claim 37,
 wherein a geometric property of the face sheet in the first circumferential zone of the acoustic liner also differs from a geometric property of the face sheet in the second circumferential zone of the acoustic liner.   
     
     
       39. The geared turbofan engine of claim 37, wherein a thickness of the face sheet of the acoustic liner in the first circumferential zone differs from a thickness of the face sheet of the acoustic liner in the second circumferential zone. 
     
     
       40. The geared turbofan engine of claim 37, wherein a porosity of the face sheet of the acoustic liner in the first circumferential zone differs from a porosity of the face sheet of the acoustic liner in the second circumferential zone. 
     
     
       41. The geared turbofan engine of claim 40, wherein a thickness of the face sheet of the acoustic liner in the first circumferential zone differs from a thickness of the face sheet of the acoustic liner in the second circumferential zone. 
     
     
       42. The geared turbofan engine of claim 37, wherein a width of the resonator chambers of the cellular core of the acoustic liner in the first circumferential zone differs from a width of the resonator chambers of the cellular core of the acoustic liner in the second circumferential zone. 
     
     
       43. The geared turbofan engine of claim 37, wherein a radial cross-sectional shape of the resonator chambers of the cellular core of the acoustic liner in the first circumferential zone differs from a radial cross-sectional shape of the resonator chambers of the cellular core of the acoustic liner in the second circumferential zone. 
     
     
       44. The geared turbo fan engine of claim 37, wherein the acoustic liner is disposed at an inlet section of the bypass flow duct such that the acoustic liner defines a surface of the inlet section of the bypass flow duct. 
     
     
       45. The geared turbofan engine of claim 44, comprising a second acoustic liner disposed at a rear section of the bypass flow duct, the second acoustic liner comprising:
 a cellular core; and   a face sheet disposed on the cellular core and defining a surface of the rear section of the bypass flow duct;   wherein a geometric property of the cellular core of the second acoustic liner, a geometric property of the face sheet of the second acoustic liner, or both varies along an axial length of the rear section of the bypass flow duct.   
     
     
       46. The geared turbofan engine of claim 45, wherein the geometric property that varies along the axial length of the rear section of the bypass flow duct comprises a porosity of the face sheet of the second acoustic liner. 
     
     
       47. The geared turbofan engine of claim 45, wherein the cellular core of the second acoustic liner comprises resonator chambers, and
 wherein the geometric property that varies along the axial length of the rear section of the bypass flow duct comprises a depth of the resonator chambers of the second acoustic liner.   
     
     
       48. The geared turbofan engine of claim 37, wherein at least a portion of the acoustic liner is configured to attenuate frequencies of less than 1000 Hz. 
     
     
       49. The geared turbofan engine of claim 37, wherein at a flight condition of the geared turbofan engine, the fan is configured to rotate at a frequency of between 200 Hz and 6000 Hz. 
     
     
       50. The geared turbofan engine of claim 37, wherein at a flight condition of the geared turbofan engine, the fan pressure ratio of the fan is between 1.25 and 1.60. 
     
     
       51. A geared turbofan engine comprising:
 an engine core comprising:
 a first rotor connected to a fan; 
 a second rotor; and 
 a gear train connecting the first rotor to the second rotor; 
   a core casing disposed circumferentially around at least a portion of the engine core;   a nacelle disposed circumferentially around at least a portion of the core casing, wherein a bypass flow duct is defined between the nacelle and the core casing; and   an acoustic liner extending at least partially around a circumference of the bypass flow duct and disposed on an inner surface of the nacelle, the acoustic liner comprising:
 a cellular core; and 
 a face sheet disposed on the cellular core and defining a surface of the bypass flow duct; 
   wherein:
 a first circumferential zone of the acoustic liner extends around a first portion of the circumference of the bypass flow duct, in which the cellular core of the acoustic liner in the first circumferential zone includes multiple, adjacent resonator chambers, 
 a second circumferential zone of the acoustic liner extends around a second portion of the circumference of the bypass flow duct adjacent the first portion of the circumference, in which the cellular core of the acoustic liner in the second circumferential zone includes multiple, adjacent resonator chambers, 
 a geometric property of the cellular core or a geometric property of the face sheet in the first circumferential zone of the acoustic liner differs from the corresponding geometric property of the cellular core or geometric property of the face sheet in the second circumferential zone of the acoustic liner, and 
 an arc length of the first circumferential zone along the circumference of the bypass flow duct is different from an arc length of the second circumferential zone along the circumference of the bypass flow duct. 
   
     
     
       52. The geared turbofan engine of claim 51, wherein the first and second circumferential zones each includes multiple subzones spaced circumferentially apart from one another. 
     
     
       53. The geared turbofan engine of claim 52, wherein there are exactly two subzones in each of the first and second circumferential zones. 
     
     
       54. The geared turbofan engine of claim 51, wherein a porosity of the face sheet of the acoustic liner in the first circumferential zone differs from a porosity of the face sheet of the acoustic liner in the second circumferential zone. 
     
     
       55. The geared turbofan engine of claim 54, wherein a depth of the resonator chambers of the cellular core in the first circumferential zone differs from a depth of the resonator chambers of the cellular core in the second circumferential zone. 
     
     
       56. The geared turbofan engine of claim 55, wherein a thickness of the face sheet of the acoustic liner in the first circumferential zone differs from a thickness of the face sheet of the acoustic liner in the second circumferential zone. 
     
     
       57. The geared turbofan engine of claim 54, wherein a thickness of the face sheet of the acoustic liner in the first circumferential zone differs from a thickness of the face sheet of the acoustic liner in the second circumferential zone. 
     
     
       58. The geared turbofan engine of claim 51, wherein a depth of the resonator chambers of the cellular core in the first circumferential zone differs from a depth of the resonator chambers of the cellular core in the second circumferential zone. 
     
     
       59. The geared turbofan engine of claim 58, wherein a thickness of the face sheet of the acoustic liner in the first circumferential zone differs from a thickness of the face sheet of the acoustic liner in the second circumferential zone. 
     
     
       60. The geared turbofan engine of claim 51, wherein a width of the resonator chambers of the cellular core in the first circumferential zone differs from a width of the resonator chambers of the cellular core in the second circumferential zone. 
     
     
       61. The geared turbofan engine of claim 51, wherein at least a portion of the acoustic liner is configured to attenuate frequencies of less than 1000 Hz. 
     
     
       62. The geared turbofan engine of claim 51, wherein at a flight condition of the geared turbofan engine, the fan is configured to rotate at a frequency of between 200 Hz and 6000 Hz. 
     
     
       63. The geared turbofan engine of claim 51, wherein at a flight condition of the geared turbofan engine, the fan pressure ratio of the fan is between 1.25 and 1.60. 
     
     
       64. The geared turbofan engine of claim 51, wherein a thickness of the face sheet of the acoustic liner in the first circumferential zone differs from a thickness of the face sheet of the acoustic liner in the second circumferential zone.