Patent Publication Number: US-9890657-B2

Title: Acoustic treatment in geared turbomachine

Description:
This application is a continuation of U.S. application Ser. No. 13/406,712, filed on Feb. 28, 2012. 
    
    
     BACKGROUND 
     This disclosure relates generally to acoustic treatments and, more particularly, to acoustic treatments that establish a portion of the core cowl aft a fan cowl. 
     Turbomachines, such as gas turbine engines, typically include a fan section, a turbine section, a compressor section, and a combustor section. Turbomachines may employ a geared architecture that allows the fan to spin at a slower rotational speed than the low pressure turbine. 
     Air moves into the turbomachine through the fan section. Some of this air moves into a core of the turbomachine. The remaining air moves through a bypass flowpath established between a fan cowl and a core engine cowl of the core. The core of the turbomachine extends axially outside the fan cowl. That is, some of the bypass stream is ducted. 
     Acoustic treatments attenuate noise radiating from the turbomachine. These acoustic treatments are traditionally limited to ducted areas of the turbomachine. Extending acoustic treatments past an aft end of the fan cowl is less efficient as a result of the reduced interaction between the propagating noise and the acoustic treatment surface. Moreover, placing the acoustic treatment on the core engine cowl aft the fan cowl subject to high Mach number flows (e.g. near a nozzle exit) results in a performance penalty associated with increased drag. In a typical turbofan with a fan pressure ratio greater than 1.5, the high performance penalty associated with the high nozzle Mach number along with the reduced noise benefit makes it impractical to place acoustic treatment on the core engine cowl past the fan nozzle exit plane. 
     SUMMARY 
     A geared turbomachine assembly according to an exemplary aspect of the present disclosure includes, among other things, an acoustic treatment that establishes a portion of the core cowl aft the fan cowl. 
     In a further non-limiting embodiment of the foregoing geared turbomachine assembly, a portion of the acoustic treatment may be axially aft a fan cowl of the geared turbomachine. 
     In a further non-limiting embodiment of either of the foregoing geared turbomachine assemblies, a portion of the acoustic treatment is axially aligned with a low-pressure turbine section of the geared turbomachine. 
     In a further non-limiting embodiment of any of the foregoing geared turbomachine assemblies, the acoustic treatment may extend continuously from the ducted area to the portion that is axially aligned with the low-pressure turbine section of the geared turbomachine. 
     In a further non-limiting embodiment of any of the foregoing geared turbomachine assemblies, the acoustic treatment may form a portion of a core engine cowl. 
     In a further non-limiting embodiment of any of the foregoing geared turbomachine assemblies, the acoustic treatment may extend axially to an aftmost end of the core engine cowl. 
     In a further non-limiting embodiment of any of the foregoing geared turbomachine assemblies, the acoustic treatment may extend continuously from the ducted area to the aftmost end. 
     In a further non-limiting embodiment of any of the foregoing geared turbomachine assemblies, the acoustic treatment may comprise a perforated face sheet having an open area density that is from 4% to 30%. 
     In a further non-limiting embodiment of any of the foregoing geared turbomachine assemblies, the acoustic treatment may comprise a microperforated face sheet having apertures with a diameter that is less than 25 percent of the perforated face sheet. 
     In a further non-limiting embodiment of any of the foregoing geared turbomachine assemblies, the geared turbomachine assembly may have a fan pressure ratio that is less than 1.45 to 1. 
     A geared turbomachine assembly according to another exemplary aspect of the present disclosure includes, among other things, a fan cowl housing a fan that is driven by a geared architecture, a core cowl housing a core engine, an acoustic treatment establishing a portion of the core cowl aft the fan cowl. 
     In a further non-limiting embodiment of the foregoing geared turbomachine assembly, the geared turbomachine assembly may have a portion of the acoustic treatment axially aft a fan cowl of the geared turbomachine. 
     In a further non-limiting embodiment of either of the foregoing geared turbomachine assemblies, a portion of the acoustic treatment may be axially aligned with a low-pressure turbine section of the geared turbomachine. 
     In a further non-limiting embodiment of any of the foregoing geared turbomachine assemblies, the acoustic treatment may form a portion of the core cowl. 
     In a further non-limiting embodiment of any of the foregoing geared turbomachine assemblies, the acoustic treatment may extend axially to an aftmost end of the core cowl. 
     In a further non-limiting embodiment of any of the foregoing geared turbomachine assemblies, the acoustic treatment may comprise a perforated face sheet having an open area density that is from 4% to 30%. 
     In a further non-limiting embodiment of any of the foregoing geared turbomachine assemblies, the acoustic treatment may comprise a microperforated face sheet having apertures with a diameter that is less than 25 percent of the perforated face sheet 
     In a further non-limiting embodiment of any of the foregoing geared turbomachine assemblies, the geared turbomachine assembly may have a fan pressure ratio that is less than 1.45 to 1. 
     A method of attenuating noise in a geared turbomachine according to another exemplary aspect of the present disclosure includes, among other things, attenuating noise axially aft of a fan cowl using an acoustic treatment that provides a portion of an engine core, wherein the fan cowl houses a fan driven by a geared architecture. 
     In a further non-limiting embodiment of the foregoing method of attenuating noise in a geared turbomachine, the acoustic treatment may extend axially past a high-pressure turbine section. 
    
    
     
       DESCRIPTION OF THE FIGURES 
       The various features and advantages of the disclosed examples will become apparent to those skilled in the art from the detailed description. The figures that accompany the detailed description can be briefly described as follows: 
         FIG. 1  shows a schematic view of an example turbomachine having a geared architecture. 
         FIG. 2  shows a cross-section view of a portion of another example turbomachine having a geared architecture. 
         FIG. 3  shows a partial section view of an example acoustic treatment used in the  FIG. 2  engine. 
         FIG. 4  shows a partial section view of another example acoustic treatment used in the  FIG. 2  engine. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  schematically illustrates an example turbomachine, which is a gas turbine engine  20  in this example. The gas turbine engine  20  is a two-spool turbofan gas turbine engine that generally includes a fan section  22 , a compressor section  24 , a combustion section  26 , and a turbine section  28 . Other examples may include an augmentor section (not shown) among other systems or features. 
     Although depicted as a two-spool turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with turbofans. That is, the teachings may be applied to other types of turbomachines and turbine engines including three-spool architectures. 
     The example engine  20  generally includes a low-speed spool  30  and a high-speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36 . The low-speed spool  30  and the high-speed spool  32  are rotatably supported by several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively, or additionally, be provided. 
     The low-speed spool  30  generally includes an inner shaft  40  that interconnects a fan  42 , a low-pressure compressor  44 , and a low-pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a geared architecture  48  to drive the fan  42  at a lower speed than the low-speed spool  30 . 
     The high-speed spool  32  includes an outer shaft  50  that interconnects a high-pressure compressor  52  and high-pressure turbine  54 . 
     The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A, which is collinear with the longitudinal axes of the inner shaft  40  and the outer shaft  50 . 
     The combustion section  26  includes a circumferentially distributed array of combustors  56  generally arranged axially between the high-pressure compressor  52  and the high-pressure turbine  54 . 
     In some non-limiting examples, the engine  20  is a high-bypass geared aircraft engine. In a further example, the engine  20  bypass ratio is greater than about six (6 to 1). 
     The geared architecture  48  of the example engine  20  includes an epicyclic gear train, such as a planetary gear system or other gear system. The example epicyclic gear train has a gear reduction ratio of greater than about 2.3 (2.3 to 1). 
     The low-pressure turbine  46  pressure ratio is pressure measured prior to inlet of low-pressure turbine  46  as related to the pressure at the outlet of the low-pressure turbine  46  prior to an exhaust nozzle of the engine  20 . In one non-limiting embodiment, the bypass ratio of the engine  20  is greater than about ten (10 to 1), the fan diameter is significantly larger than that of the low pressure compressor  44 , and the low-pressure turbine  46  has a pressure ratio that is greater than about 5 (5 to 1). The geared architecture  48  of this embodiment is an epicyclic gear train with a gear reduction ratio of greater than about 2.5 (2.5 to 1). It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans. 
     In this embodiment of the example engine  20 , a significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. 
     The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. This flight condition, with the engine  20  at its best fuel consumption, is also known as “Bucket Cruise” Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust. 
     Fan Pressure Ratio is the pressure ratio across a blade of the fan section  22  without the use of a Fan Exit Guide Vane system. The Fan Pressure Ratio according to one non-limiting embodiment of the example engine  20  is less than 1.45 (1.45 to 1), which is relatively low. 
     Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of Temperature divided by 518.7^0.5. The Temperature represents the ambient temperature in degrees Rankine. The Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example engine  20  is less than about 1150 fps (351 m/s). 
     The example gas turbine engine  20  generally includes an area  60  and an area  62 . In this example, the area  60  represents the portions of the gas turbine engine  20  axially aligned with, and radially bounded by, a fan cowl  66  (or nacelle). The area  62  represents the portions of the gas turbine engine  20  axially outside of the fan cowl  66 . 
     In this example, a core engine cowl  70  of the gas turbine engine  20  includes portions in the area  60  and portions in the area  62 . The core engine cowl  70  includes an acoustic treatment  74  that extends continuously from the area  60  to the area  62 . The acoustic treatment  74  is on an outer surface of the core engine cowl  70 . That is, at least some of the acoustic treatment  74  is in an area of the gas turbine engine  20  axially aft the fan cowl  66 . 
     Referring to  FIG. 2 , another example geared turbomachine is the gas turbine engine  120 . In this example, the acoustic treatment  74  extends from the area  160 , past an aft end  76  of the fan cowl  166 , to the area  162  of the engine  120 . Notably, the example acoustic treatment  74  extends continuously, and without interruption, from the area  160  to the aftmost end  84  of the core engine cowl  170 . 
     A portion of the acoustic treatment  74  is axially aligned with the turbine section  128  of the engine  120 , and, in particular, the low-pressure turbine  146  within the turbine section  128 . The acoustic treatment  74  is continuous and uninterrupted from the area  160 , past the high-pressure turbine  154 , to the portion that is aligned with the low-pressure turbine  146 . In this example, the acoustic treatment  74  also lines radially inwardly facing surfaces of the fan cowl  166 . 
     Referring to  FIG. 3  with continuing reference to  FIG. 2 , an example acoustic treatment  74  includes a perforated face sheet  78  and a backing sheet  82 . A honeycomb structure  86  is sandwiched between the perforated face sheet  78  and the backing sheet  82 . The perforated face sheet  78 , the backing sheet  82  and the honeycomb structure are all aluminum in this example. Other examples of the acoustic treatment  74  may be made of a composite material or some other material. 
     The example perforated face sheet  78  includes an array of perforations  88  each having a diameter that is from 0.03 inches (0.762 mm) to 0.06 inches (1.524 mm). A density of the open area in the perforated face sheet  78  is from 6% to 20% in this example. Other perforated face sheets  78  have perforation  88  having other diameters, such as 4% to 30%. 
     In this example, the core engine cowl  170  in the areas of the acoustic treatment  74  is formed entirely by the acoustic treatment  74 . That is, there is no walled structure that the acoustic treatment  74  is secured to. The acoustic treatment  74  is at an outer surface of the core engine cowl  170 . 
     In another example, the acoustic treatment  74  and, in particular, the back sheet of the acoustic treatment  74 , is adhesively secured to a wall (not shown). In such an example, the adhesively secured acoustic treatment and the wall together form the core engine cowl  170 . 
     The acoustic treatment  74  is a perforated acoustic treatment with a single degree of freedom (SDOF) in this example, other examples may include a Double Degree Of Freedom (DDOF) liner, bulk absorber liner or a microperforated acoustic treatment  174  as shown in  FIG. 4 . The microperforated acoustic treatment  174  includes a perforated face sheet  178  and a backing sheet  182  sandwiching a honeycomb structure  186 . Perforations  188  in the perforated face sheet  178  each have a diameter that is less than 25 percent of a thickness t of the perforated face sheet  178 . 
     Features of the disclosed examples include incorporating an acoustic treatment in an area of a turbomachine that is traditionally not well suited for acoustic treatments due to the associated drag penalties. The added acoustic treatment has minimal impact on fuel burn. 
     In this disclosure, like reference numerals designate like elements where appropriate, and reference numerals with the addition of one-hundred or multiples thereof designate modified elements. The modified elements incorporate the same features and benefits of the corresponding modified elements, expect where stated otherwise. 
     The preceding description is exemplary rather than limiting in nature. Variations and modifications to the disclosed examples may become apparent to those skilled in the art that do not necessarily depart from the essence of this disclosure. Thus, the scope of legal protection given to this disclosure can only be determined by studying the following claims.