Abstract:
A nacelle assembly according to an exemplary aspect of the present disclosure includes, among other things, a core nacelle defined at least partially about a core engine, a fan nacelle mounted at least partially around the core nacelle to define a fan bypass flow path, and a variable area fan nozzle in communication with the fan bypass flow path. The variable area fan nozzle has a first fan nacelle section and a second fan nacelle section downstream of the first fan nacelle section. The first fan nacelle section and the second fan nacelle section are axially movable relative to one another to define an auxiliary port to vary a fan nozzle exit area and adjust fan bypass airflow. The auxiliary port is defined between the first fan nacelle section and the second fan nacelle section. The first fan nacelle section comprises a first acoustic system which provides an acoustic impedance configured to attenuate a noise characterized by a leading edge of the second fan nacelle section. The first acoustic system is defined at least in part within a trailing edge region of the first fan nacelle section. A method of reducing a total effective perceived noise level of a gas turbine engine with a variable area fan nozzle is also disclosed.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application is a continuation of U.S. application Ser. No. 12/804,666, filed Jul. 27, 2010 and published as U.S. Publication No. 2012/0023901. 
         [0002]    Reference is made to co-pending commonly-assigned, U.S. patent application Pub. No.: US 2009/0320488, entitled “GAS TURBINE ENGINE WITH NOISE ATTENUATING VARIABLE AREA FAN NOZZLE,” and co-pending commonly-assigned, U.S. patent application (application Ser. No. 12/838,620, UTC docket No. PA-0013449-US), entitled “GAS TURBINE ENGINE WITH NOISE ATTENUATING VARIABLE AREA FAN NOZZLE”. 
     
    
     BACKGROUND 
       [0003]    The present invention relates to a gas turbine engine, and particularly to a bypass turbofan engine nacelle assembly, and more particularly to a bypass turbofan engine nacelle assembly having a ported variable area fan nozzle (VAFN) with an acoustic system to reduce the total effective perceived noise level (EPNL). 
         [0004]    Gas turbine engines which have an engine cycle modulated with a VAFN typically provide a smaller fan exit nozzle area during cruise conditions and a larger fan exit nozzle area during take-off and landing conditions. 
         [0005]    The ported VAFN system may generate significant noise as upstream turbulence interacts with the leading edge of the VAFN. The upstream turbulence may result from turbulent boundary layers which expand from the upstream fixed nacelle wall, turbulence which evolves from the upstream fan exit guide vane (FEGV) wakes or endwall effects, and flow separation that may occur from the contour of the upstream nacelle wall. The physical mechanism for leading edge VAFN noise, which exhibits acoustic dipole behavior, is fundamentally different from traditional jet exhaust mixing noise, which exhibits acoustic quadrupole behavior. Additionally, this excess noise is not significantly reduced in forward flight as typical jet exhaust mixing noise. Thus, VAFN noise contributes toward the total effective perceived noise level. Accordingly, it is desirable to reduce VAFN noise. 
       SUMMARY 
       [0006]    A nacelle assembly for a gas turbine engine according to an example of the present disclosure includes a core nacelle defined about an engine centerline axis, a fan nacelle mounted at least partially around the core nacelle, the core nacelle and fan nacelle defining a fan bypass flow path, and a variable area fan nozzle in communication with the fan bypass flow path. The variable area fan nozzle has a first fan nacelle section and a second fan nacelle section downstream of the first fan nacelle section. The first fan nacelle section and the second fan nacelle section axially move relative to one another to define an auxiliary port to vary a fan nozzle exit area and adjust fan bypass airflow. The auxiliary port is defined between the first fan nacelle section and the second fan nacelle section. The first fan nacelle section comprises a first acoustic system having a first acoustic impedance configured to attenuate a noise characterized by a leading edge of the second fan nacelle section. The first acoustic system is defined at least in part within a trailing edge region of the first fan nacelle section. 
         [0007]    In a further embodiment of any of the foregoing embodiments, the first acoustic system extends axially from a trailing edge of the first fan nacelle section. 
         [0008]    In a further embodiment of any of the foregoing embodiments, the first acoustic system comprises a perforated sheet. 
         [0009]    In a further embodiment of any of the foregoing embodiments, the perforated sheet further comprises a deformable material. 
         [0010]    In a further embodiment of any of the foregoing embodiments, the first acoustic system comprises a perforated sheet, a backing plate and internal partitions. 
         [0011]    In a further embodiment of any of the foregoing embodiments, the first acoustic system comprises a bulk acoustic absorbing material. 
         [0012]    In a further embodiment of any of the foregoing embodiments, the first acoustic system comprises a bulk cartridge. 
         [0013]    In a further embodiment of any of the foregoing embodiments, the bulk cartridge comprises a bulk acoustic absorbing material. 
         [0014]    In a further embodiment of any of the foregoing embodiments, the bulk absorbing material is selected from a group consisting of sintered metal, ceramic foam, aramid fiber, carbide material and composite. 
         [0015]    In a further embodiment of any of the foregoing embodiments, the bulk absorbing material comprises a deformable material. 
         [0016]    In a further embodiment of any of the foregoing embodiments, the second fan nacelle section comprises a second acoustic system having a second acoustic impedance configured to attenuate a noise characterized by the leading edge of the second fan nacelle section. 
         [0017]    In a further embodiment of any of the foregoing embodiments, the second acoustic system is defined at least in part within a leading edge region of the second fan nacelle section. 
         [0018]    In a further embodiment of any of the foregoing embodiments, the first acoustic system and the second acoustic system are placed in close proximity to form a region of opposed acoustic treatment along the auxiliary port. 
         [0019]    In a further embodiment of any of the foregoing embodiments, the first acoustic system extends axially a first length along the first nacelle section and the second acoustic system extends axially a second length along the second nacelle section. The first length and the second length is substantially equal to each other. 
         [0020]    In a further embodiment of any of the foregoing embodiments, the second fan nacelle section defines a third length between a leading edge and a trailing edge of the second fan nacelle section. The second length is a ratio of between 1:4 and 1:3 of the third length. 
         [0021]    In a further embodiment of any of the foregoing embodiments, the second fan nacelle section is subdivided into a multiple of independently operable sectors. Each of the multiple of independently operable sectors axially movable relative the first fan nacelle section to define an asymmetric fan nozzle exit area. 
         [0022]    In a further embodiment of any of the foregoing embodiments, the first acoustic system is configured to extend between a trailing edge of the first fan nacelle section and a point defined on the first fan nacelle section at which a leading edge of the second fan nacelle section engages the first fan nacelle section when the variable area fan nozzle is located in a closed position. 
         [0023]    A bypass gas turbine engine according to an example of the present disclosure includes a core engine defined about an axis, a core nacelle defined at least partially about the core engine, a fan nacelle mounted at least partially around the core nacelle to define a fan bypass flow path, and a variable area fan nozzle in communication with the fan bypass flow path, the variable area fan nozzle having a first fan nacelle section and a second fan nacelle section downstream of the first fan nacelle section. The first fan nacelle section and the second fan nacelle section axially movable relative to one another to define an auxiliary port to vary a fan nozzle exit area and adjust fan bypass airflow. The auxiliary port is defined between the first fan nacelle section and the second fan nacelle section. The first fan nacelle section comprises a first acoustic system which provides an acoustic impedance configured to attenuate a noise characterized by a leading edge of the second fan nacelle section. The first acoustic system is defined at least in part within a trailing edge region of the first fan nacelle section. 
         [0024]    In a further embodiment of any of the foregoing embodiments, the second fan nacelle section comprises a second acoustic system having a second acoustic impedance configured to attenuate a noise characterized by the leading edge of the second fan nacelle section. 
         [0025]    In a further embodiment of any of the foregoing embodiments, the first acoustic system and the second acoustic system are placed in close proximity to form a region of opposed acoustic treatment along the auxiliary port. 
         [0026]    In a further embodiment of any of the foregoing embodiments, the first acoustic system extends axially from a trailing edge of the first fan nacelle section. 
         [0027]    A further embodiment of any of the foregoing embodiments includes a gear train driven by the core engine, and a turbofan driven by the gear train about the axis. 
         [0028]    A method of reducing a total effective perceived noise level of a gas turbine engine with a variable area fan nozzle according to an example of the present disclosure includes axially moving a second fan nacelle section between a closed position in which the second fan nacelle section is in sequential alignment with a first fan nacelle section in response to a cruise flight condition and an open positioning in which the second fan nacelle section is aftward of the first fan nacelle section to define an auxiliary port in response to a non-cruise flight condition. The auxiliary port is defined between the first fan nacelle section and the second fan nacelle section. The first fan nacelle section has a trailing edge region with a first acoustic system comprising a first acoustic absorbing material which provides an acoustic impedance configured to attenuate a noise characterized by a leading edge of the second fan nacelle section when the second fan nacelle section is positioned at a non-closed position. 
         [0029]    A further embodiment of any of the foregoing embodiments includes a second acoustic system having a second acoustic impedance and forming at least a portion of leading edge region of the second fan nacelle section. The second acoustic impedance configured to attenuate a noise characterized by the leading edge of the second fan nacelle section. 
         [0030]    A further embodiment of any of the foregoing embodiments includes placing the first acoustic system and the second acoustic system in close proximity to form a region of opposed acoustic treatment along the auxiliary port. 
         [0031]    In a further embodiment of any of the foregoing embodiments, the first acoustic system extends axially from a trailing edge of the first fan nacelle section. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0032]    The various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description of the currently preferred embodiments. The drawings that accompany the detailed description can be briefly described as follows: 
           [0033]      FIG. 1  is a general schematic partial fragmentary view of an exemplary gas turbine engine embodiment for use with the present invention; 
           [0034]      FIG. 2  is a rear view of the engine; 
           [0035]      FIG. 3A  is a side view of the engine integrated with a pylon; 
           [0036]      FIG. 3B  is a rear perspective view of the engine integrated with a pylon; 
           [0037]      FIG. 4A  is a sectional side view of the ported VAFN in a closed position; 
           [0038]      FIG. 4B  is a sectional side view of the ported VAFN in an open position; 
           [0039]      FIG. 5  is a sectional side view of the ported VAFN with an acoustic system; 
           [0040]      FIG. 6A  is a sectional side view of the acoustic system; 
           [0041]      FIG. 6B  is a sectional side view of the acoustic system; 
           [0042]      FIG. 6C  is a sectional side view of the acoustic system; 
           [0043]      FIG. 7  is a schematic sectional side view of the ported VAFN with a tandem acoustic system; 
           [0044]      FIG. 8A  is a sectional side view of the acoustic system; 
           [0045]      FIG. 8B  is a sectional side view of the acoustic system; and 
           [0046]      FIG. 8C  is a sectional side view of the acoustic system 
       
    
    
     DETAILED DESCRIPTION 
       [0047]      FIG. 1  illustrates a general partial fragmentary schematic view of a gas turbofan engine  10 , which is circumferentially disposed about an engine axis of rotation A, suspended from an engine pylon P within an engine nacelle assembly N. 
         [0048]    The turbofan engine  10  includes a core engine within a core nacelle  12  that houses a low spool  14  and high spool  24 . The low spool  14  includes a low pressure compressor  16  and low pressure turbine  18 . The low spool  14  also drives a fan section  20  directly or through a gear train  22 . The high spool  24  includes a high pressure compressor  26  and high pressure turbine  28 . A combustor  30  is arranged between the high pressure compressor  26  and high pressure turbine  28 . The low and high spools  14 ,  24  rotate about an engine axis of rotation A. 
         [0049]    Airflow enters a fan nacelle  34 , which at least partially surrounds the core nacelle  12 . The fan section  20  communicates airflow into the core nacelle  12 , one portion of the air (referred to as bypass airflow, hereinafter) flows through the bypass flow path  40 , and the other portion of air (referred to as core airflow, hereinafter) flows through the core flow path  41 . Core airflow compressed by the low pressure compressor  16  and the high pressure compressor  26  is mixed with the fuel and burned in the combustor  30  and expanded over the high pressure turbine  28  and low pressure turbine  18  sequentially. The turbines  28 ,  18  are coupled for rotation with, respective, spools  24 ,  14  to rotationally drive the compressors  26 ,  16  and through the gear train  22 , the fan section  20  in response to the expansion. A core engine exhaust E exits the core nacelle  12  through a core nozzle  43  defined between the core nacelle  12  and a tail cone  32 . 
         [0050]    The core nacelle  12  is supported within the fan nacelle  34  by circumferentially space structures  36  often generically referred to as Fan Exit Guide Vanes (FEGV). The bypass flow path  40  is defined between the core nacelle  12  and the fan nacelle  34 . The engine  10  generates a bypass flow arrangement with a bypass ratio in which substantial percent of the airflow which enters the fan nacelle  34  becomes bypass flow B. The bypass flow B communicates through the generally annular bypass flow path  40  and is discharged from the engine  10  through a ported variable area fan nozzle  42  which defines a nozzle exit area  44  between the fan nacelle  34  and the core nacelle  12  at a fan nacelle end segment  34 S of the fan nacelle  34  downstream of the fan section  20 . 
         [0051]    The ported VAFN  42  (referred to as VAFN, hereinafter) operates to effectively vary the area of the fan nozzle exit area  44  to selectively adjust the pressure ratio of the bypass flow B in response to a controller C. Low fan pressure ratio turbofans are desirable for their high propulsive efficiency. However, low fan pressure ratio fans may be relatively susceptible to fan stability/flutter problems at low power and low flight speeds. The VAFN allows the engine to change to a more favorable fan operating line at low power, avoiding any instability region, and still provide the relatively smaller nozzle area necessary to obtain a high-efficiency fan operating line at cruise. 
         [0052]    A significant amount of thrust is provided by the bypass flow B. The fan section  20  of the engine  10  is preferably designed for a particular flight condition—for example cruise at about 0.8 Mach number and 35,000 feet. As the fan blades within the fan section  20  are typically (but not necessarily) fixed and designed at a particular stagger angle for an efficient cruise condition, the VAFN  42  may be operated to effectively vary the fan nozzle exit area  44  to adjust fan bypass air flow such that the angle of attack or incidence on the fan blades is maintained close to the design incidence for efficient engine operation at other flight conditions, such as landing and takeoff to thus provide optimized engine operation over a range of flight conditions with respect to performance and other performance parameters such as noise levels. 
         [0053]      FIG. 1  shows a two-spool, geared architecture turbofan engine  10 , it is understood, however, that the disclosure is applicable to any gas turbine engine having a ported variable area fan nozzle. Those gas turbine engine architectures include, but not limited to, three-spool, two-spool or single-spool, geared architecture or direct drive turbofan gas turbine engines. 
         [0054]    The VAFN  42 , in one exemplary embodiment, can be extended circumferentially over 360° about the engine axis of rotation A; and in another exemplary embodiment, may be separated into at least two sectors  42 A- 42 B ( FIG. 2 ) defined between the pylon P and a lower Bi-Fi splitter L. It should be understood that although two segments are illustrated, any number of sectors and segments may alternatively or additionally be provided. 
         [0055]    The VAFN  42  generally includes an auxiliary port system  50  having a first fan nacelle section  52  and a second fan nacelle section  54  movably mounted relative the first fan nacelle section  52 . The second fan nacelle section  54  axially slides along the engine axis A relative the fixed first fan nacelle section  52  to change the effective area of the fan nozzle exit area  44 . The second fan nacelle section  54 , in one non-limiting embodiment, slides aftward upon a track fairing  56 A,  56 B (illustrated schematically in  FIGS. 3A and 3B ) in response to an actuator  58  (illustrated schematically in  FIG. 1 ). The track fairing  56 A,  56 B extends from the first fan nacelle section  52  adjacent the respective pylon P and the lower Bi-Fi splitter L ( FIG. 3B ). 
         [0056]    The VAFN  42  changes the physical area and geometry of the bypass flow path  40  during particular flight conditions. The bypass flow B is effectively altered by sliding of the second fan nacelle section  54  relative the first fan nacelle section  52  between a closed position ( FIG. 4A ) and an open position ( FIG. 4B ). The auxiliary port system  50  is closed by positioning the second fan nacelle section  54  in-line with the first fan nacelle section  52  to define the fan nozzle exit area  44  as exit area F0 ( FIG. 4A ). The VAFN  42  is opened by moving the second fan nacelle section  54  aftward along the track fairing  56 A,  56 B away from the first fan nacelle section  52  to open the auxiliary port  60  ( FIG. 4B ) which extends between the second fan nacelle section  54  relative the first fan nacelle section  52 , and provide an increased fan nozzle exit area  44  exit area F1. That is, the exit area F1 with the auxiliary port  60  ( FIG. 4B ) is greater than exit area F0 ( FIG. 4A ), thanks to the cone-shaped core nacelle  12 . 
         [0057]    In one non-limiting embodiment, the auxiliary port  60  is incorporated within the bypass flow path  40  aft of the Fan Exit Guide Vanes  36  (FEGVs). The auxiliary port  60  is located through the bypass duct outer wall. 
         [0058]    In operation, the VAFN  42  communicates with the controller C to move the second fan nacelle section  54  relative the first fan nacelle section  52  of the auxiliary port system  50  to effectively vary the area defined by the fan nozzle exit area  44 . Various control systems including an engine controller or an aircraft flight control system may also be usable with the present invention. By adjusting the axial position of the entire periphery of the second fan nacelle section  54  in which all sectors are moved simultaneously, engine thrust and fuel economy are optimized during each flight regime by varying the fan nozzle exit area. 
         [0059]    The VAFN may generate significant noise as upstream airflow turbulence interacts with the leading edge  62   e  of the second fan nacelle section  54 . To reduce the VAFN noise, referring to  FIG. 5 , the first fan nacelle section  52  includes a trailing edge region  140  having a first acoustic system  142 . The acoustic system  142  utilizes the available volume of the trailing edge region  140  to achieve optimal acoustic impedance. The physical size and location for the trailing edge region  140  depend on the specific VAFN configuration. A rule of thumb for the size of the trailing edge region  140  is that it should extend in an upstream direction from the trailing edge  140   e  of the first fan nacelle section  52  to the point where the leading edge  62   e  of the second fan nacelle section  54  makes contact with the first fan nacelle section  52  when the second fan nacelle section  54  is positioned at closed position. With acoustic impedance, the acoustic system  142  operates to attenuate the leading edge region  62  noise. As a result, far field total effective perceived noise level is reduced. Since the acoustic system  142  is naturally placed in close proximity to the VAFN noise source at the VAFN leading edge, optimal noise reduction is achieved. 
         [0060]    Referring to  FIG. 6A , the acoustic system  142  includes a perforated sheet  180 , a backing plate  182  and internal partitions  184 . The perforated sheet  180  covers inner surface of the trailing edge region  140 ; the backing plate  182  defines characteristic depth associated with acoustic frequency tuning; and the internal partitions  184  act to ensure locally reacting characteristics of the acoustic treatment. That is, the acoustic system  142  is made to be locally reacting when it contains internal partitions  184  which subdivide the inner volume of the acoustic system  142 . The internal partitions  184  which are substantially perpendicular to the perforated sheet  180  enables the acoustic waves to see a cavity depth that is equivalent to the distance between the perforated sheet  180  and the backing plate  182 . The porosity of the perforated sheet  180 , the placement of the internal partitions  184  and the backing plate  182  are designed to tune the acoustic system to achieve maximum acoustic attenuation. It should be understood that the internal partition  184  and the backing plate  182  are illustrated in partial schematic cross-section and that various arrangements of the structures may be provided to support the perforated sheet  180 . For example, in addition to the internal partitions  184  being placed in the region of the cross section shown in  FIG. 6A , the partitions may also be placed circumferentially. The pore shape, size and spacing between pores of the perforated sheet  180  could be varied to achieve optimal acoustic impedance. One of ordinary skill in the art will be able to readily determine the specific shape, size, spacing and various arrangements of the structures of the example embodiment of the present invention. 
         [0061]    Referring to  FIG. 6B , the acoustic system  142  includes a bulk acoustic absorbing material  80  such as, for example only, a sintered metal, a ceramic foam, aramid fibers under the trade name KEVLAR (a registered trademark of E.I. DuPont de Nemours &amp; Company), a composite or a carbide material, to minimize effects on the steady flow through the auxiliary port  60  and maximize effects on unsteady acoustic loading. As with the embodiment illustrated in  FIG. 6A , the porosity, and material characteristics of the bulk absorbing material  80  are selected for optimal acoustic impedance and thus optimal acoustic attenuation.  FIG. 6B  shows a completely acoustic absorbing-material filled volume in the trailing edge region  140 ; while  FIG. 6C  delineates an acoustic system  142  having a bulk cartridge  146  stuffed with acoustic absorbing material in the trailing edge region  140 . The inner surface  144  of the bulk cartridge  146  can be made from a perforated sheet or perforated elastomeric face plate. The acoustic absorbing material illustrated in both  FIGS. 6B and 6C  can also be deformable material such as, for example only, elastomeric material with mechanical properties selected to provide sound attenuation. The deformable acoustic system can be spanned larger or smaller along the inner surface  144  of the trailing edge region  140 . Additional benefit of the elastomeric material for the acoustic system is that it deforms when the VAFN is positioned in the closed position, thus serving the function of sealing the gap between the first fan nacelle section  52  and the second fan nacelle section  54  that may exist otherwise, and preventing air flow leakage through the gap. Air flow leakage through the gap may compromise gas turbine engine propulsion efficiency and operability when the VAFN is positioned in the closed position. 
         [0062]    Referring to  FIG. 7 , the acoustic impedance feature includes a combination of a trailing edge region  140  with a first acoustic system  142 , and a leading edge region  62  with a second acoustic system  64 . (Hereinafter the combined acoustic impedance feature of the two acoustic systems  142  and  64  is referred to as a tandem acoustic system  150 .) In the tandem acoustic system  150 , the second acoustic system  64  of the leading edge region  62  is in close proximity to the trailing edge region  140 . As the acoustic noise source results from turbulent boundary layers along the trailing edge, and/or any other turbulence generation mechanisms along the trailing edge impinging on the leading edge  62   e  of the second fan nacelle section  54 , the second acoustic system  64  of the leading edge region  62  operates to minimize the acoustic source radiation from the leading edge region  62  and/or attenuate the leading edge noise, the first acoustic system  142  operates to further attenuate the noise from the leading edge region  62  noise. The tandem acoustic system  150  is designed with specific physical properties to capture additional noise attenuation, that is, the tandem system  150  is better than the sum of the individual acoustic system  142  and  64 . 
         [0063]    The physical size of the leading edge region  62  may extend in the downstream direction from the leading edge  62   e  by roughly the same axial distance as the trailing edge region  140  extends upstream direction. Or simply put, axial length L1 of the leading edge region  62  and axial length L2 of the trailing edge region  140  should be substantially equal, with a length of one quarter (¼) to one third (⅓) of the second fan nacelle section  54 . The shape of the leading edge  62   e  can also further impact noise attenuation. Co-pending commonly-assigned, U.S. patent application (application Ser. No. 12/838,620, UTC docket No. PA-0013449-US), entitled “GAS TURBINE ENGINE WITH NOISE ATTENUATING VARIABLE AREA FAN NOZZLE” discloses optimal shapes of the leading edge  62   e  for noise attenuation. 
         [0064]    The second acoustic system  64  of the leading edge region  62  is disclosed in the commonly-assigned, U.S. patent application Pub. No. US 2009/0320488, entitled “GAS TURBINE ENGINE WITH NOISE ATTENUATING VARIABLE AREA FAN NOZZLE,” The entire disclosure of the prior application is hereby incorporated by reference herein in its entirety. The acoustic system  64  of the leading edge region  62  of the tandem acoustic system  150  can be, but not limited to, any of the exemplary embodiments disclosed in U.S. patent application Pub. No. US 2009/0320488: 
         [0065]    LE1—A perforated inner face sheet, a perforated outer face sheet supported by a structure as with its  FIG. 4  embodiment; 
         [0066]    LE2—A bulk absorbing material as with its  FIG. 5  embodiment; and, 
         [0067]    LE3—A forward acoustic system and an aft acoustic system as with its  FIG. 6  embodiment. 
         [0068]    For the tandem acoustic system  150 , the second acoustic system  64  of the leading edge region  62  can also be: 
         [0069]    LE4—The second acoustic system  64  may occupy only a portion of the leading edge region  62  of the second fan nacelle section  54 . 
         [0070]      FIG. 8A , as an example, shows the second acoustic system  64  occupies only the middle portion of the leading edge region  62 . As yet another example,  FIG. 8B  shows the acoustic system occupies the forward portion of the leading edge region  62 . It is understood that the second acoustic system  64  occupies only a portion of the leading edge region  62 , and may be either one or combination of acoustic designs denoted by LE1, LE2 or LE3. It is preferred that the second acoustic system  64  of the leading edge region  62  is placed in close proximity to the trailing edge region  140  of the first fan nacelle section  52  so that a region of opposed acoustic treatment  155  is realized. 
         [0071]    The acoustic system  142  of the trailing edge region  140  of the tandem acoustic system  150  can be, but is not limited to, any one of the exemplary embodiments listed below: 
         [0072]    TE1—A perforated sheet, a backing plate and internal partitions design as with  FIG. 6A  disclosed above; 
         [0073]    TE2—A bulk acoustic absorbing material design as with  FIG. 6B  disclosed above; and, 
         [0074]    TE3—A bulk cartridge stuffed with acoustic absorbing material design as with  FIG. 6C  disclosed above. 
         [0075]    As with the second acoustic system  64  of the leading edge region  62 , the first acoustic system  142  may occupy only a portion of the trailing edge region  140 . It is preferred that the first acoustic system  142  of the trailing edge region  140  exists in close proximity to the leading edge region  62  of the second fan nacelle section  54  to accommodate the required VAFN schedule, so that a region of opposed acoustic treatment  155  may be realized as exemplarily shown in  FIG. 8C . 
         [0076]    Noise reduction on the order of approximately 3 EPNdB or greater cumulative over the certification conditions described in Federal Acquisition Regulation (FAR)  36  may be readily achieved by the acoustic system disclosed herein and include both tone and broadband reductions. 
         [0077]    The foregoing description is exemplary rather than defined by the limitations within. Many modifications and variations are possible in light of the above teachings. The preferred embodiments have been disclosed, however, one of ordinary skill in the art would recognize that certain modifications would come within the scope of this invention. It is, therefore, to be understood that within the scope of the appended claims, the invention may be practiced otherwise than as specifically described.