Abstract:
A method of designing a turbofan engine according to an exemplary aspect of the present disclosure includes, among other things, providing a fan section including a plurality of fan blades, providing a low pressure turbine driving the plurality of fan blades through a gear train, providing a fan nacelle and a core nacelle, the fan nacelle at least partially surrounding the core nacelle, providing a fan bypass flow path defined between the core nacelle and the fan nacelle, and providing a fan variable area nozzle in communication with the fan bypass flow path and defining a fan nozzle exit area between the fan nacelle and the core nacelle.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    The present disclosure is a continuation of U.S. patent application Ser. No. 15/360,092, filed Nov. 23, 2016, which is a continuation of U.S. patent application Ser. No. 13/314,365, filed Dec. 8, 2011, which is a continuation in part of U.S. patent application Ser. No. 11/843,675, filed Aug. 23, 2007 and issued as U.S. Pat. No. 8,074,440. 
     
    
     BACKGROUND 
       [0002]    The present invention relates to a gas turbine engine, and more particularly to a turbofan engine having a fan variable area nozzle (VAFN) which moves axially to change a bypass flow path area thereof. 
         [0003]    Conventional gas turbine engines generally include a fan section and a core engine with the fan section having a larger diameter than that of the core engine. The fan section and the core engine are disposed about a longitudinal axis and are enclosed within an engine nacelle assembly. 
         [0004]    Combustion gases are discharged from the core engine through a core exhaust nozzle while an annular fan flow, disposed radially outward of the primary airflow path, is discharged through an annular fan exhaust nozzle defined between a fan nacelle and a core nacelle. A majority of thrust is produced by the pressurized fan air discharged through the fan exhaust nozzle, the remaining thrust being provided from the combustion gases discharged through the core exhaust nozzle. 
         [0005]    The fan nozzles of conventional gas turbine engines have a fixed geometry. The fixed geometry fan nozzles are a compromise suitable for take-off and landing conditions as well as for cruise conditions. Some gas turbine engines have implemented fan variable area nozzles. The fan variable area nozzle provide a smaller fan exit nozzle diameter during cruise conditions and a larger fan exit nozzle diameter during take-off and landing conditions. Existing fan variable area nozzles typically utilize relatively complex mechanisms that increase overall engine weight to the extent that the increased fuel efficiency therefrom may be negated. 
       SUMMARY 
       [0006]    A turbofan engine according to the present invention includes a fan variable area nozzle (VAFN) having a first fan nacelle section and a second fan nacelle section movably mounted relative the first fan nacelle section. The second fan nacelle section axially slides relative the fixed first fan nacelle section to change the effective area of the fan nozzle exit area. The VAFN changes the physical area and geometry of the bypass flow path during particular flight conditions. The VAFN is closed by positioning the second fan nacelle section in-line with the first fan nacelle section to define the fan nozzle exit area and is opened by moving the second fan nacelle section aftward to provide an increased fan nozzle exit area. 
         [0007]    In operation, the VAFN communicates with the controller to effectively vary the area defined by the fan nozzle exit area. By adjusting the entire periphery of the second fan nacelle section in which all sectors are moved simultaneously, engine thrust and fuel economy are maximized during each flight regime by varying the fan nozzle exit area. By separately adjusting circumferential sectors of the second fan nacelle section to provide an asymmetrical fan nozzle exit area, engine bypass flow is selectively vectored to provide, for example only, trim balance, thrust controlled maneuvering, enhanced ground operations and short field performance. 
         [0008]    The present invention therefore provides an effective, lightweight fan variable area nozzle for a gas turbine engine. 
         [0009]    A nacelle assembly for a high-bypass gas turbine engine according to an exemplary aspect of the present disclosure may include a core nacelle defined about an engine centerline axis, a fan nacelle mounted at least partially around the core nacelle to define a fan bypass flow path for a fan bypass airflow, and a fan variable area nozzle axially movable relative the fan nacelle to define an auxiliary port to vary a fan nozzle exit area and adjust a pressure ratio of the fan bypass airflow during engine operation. 
         [0010]    In a further non-limiting embodiment of any of the foregoing nacelle assembly embodiments, the controller may be operable to control the fan variable area nozzle to vary a fan nozzle exit area and adjust the pressure ratio of the fan bypass airflow. 
         [0011]    In a further non-limiting embodiment of any of the foregoing nacelle assembly embodiments, the controller may be operable to reduce the fan nozzle exit area at a cruise flight condition. 
         [0012]    In a further non-limiting embodiment of any of the foregoing nacelle assembly embodiments, the controller may be operable to control the aid fan nozzle exit area to reduce a fan instability. 
         [0013]    In a further non-limiting embodiment of any of the foregoing nacelle assembly embodiments, the fan variable area nozzle may define a trailing edge of the fan nacelle. 
         [0014]    In a further non-limiting embodiment of any of the foregoing nacelle assembly embodiments, the assembly may further include a controller operable to axially move the fan variable area nozzle to vary the fan nozzle exit area in response to a flight condition. 
         [0015]    In a further non-limiting embodiment of any of the foregoing nacelle assembly embodiments, the fan variable area nozzle may be aligned with the fan nacelle to define a closed position of the fan nozzle exit area. Additionally or alternatively, the fan variable area nozzle is axially offset from the fan nacelle to define an open position of the fan nozzle exit area. 
         [0016]    In a further non-limiting embodiment of any of the foregoing nacelle assembly embodiments, the nacelle assembly may further include a gear system driven by the core engine within the core nacelle to drive the fan within the fan nacelle, the gear system defines a gear reduction ratio of greater than or equal to about 2.3. 
         [0017]    In a further non-limiting embodiment of any of the foregoing nacelle assembly embodiments, the nacelle assembly may further include a gear system driven by the core engine within the core nacelle to drive the fan within the fan nacelle, the gear system defines a gear reduction ratio of greater than or equal to about 2.5. 
         [0018]    In a further non-limiting embodiment of any of the foregoing nacelle assembly embodiments, the nacelle assembly may further include a gear system driven by the core engine to drive the fan, the gear system defines a gear reduction ratio of greater than or equal to 2.5. 
         [0019]    In a further non-limiting embodiment of any of the foregoing nacelle assembly embodiments, the core engine may include a low pressure turbine which defines a pressure ratio that is greater than about five (5). 
         [0020]    In a further non-limiting embodiment of any of the foregoing nacelle assembly embodiments, the core engine may include a low pressure turbine which defines a pressure ratio that is greater than five (5). 
         [0021]    In a further non-limiting embodiment of any of the foregoing nacelle assembly embodiments, the bypass flow may define a bypass ratio greater than about six (6). 
         [0022]    In a further non-limiting embodiment of any of the foregoing nacelle assembly embodiments, the bypass flow may define a bypass ratio greater than about ten (10). 
         [0023]    In a further non-limiting embodiment of any of the foregoing nacelle assembly embodiments, the bypass flow may define a bypass ratio greater than ten (10). 
         [0024]    A gas turbine engine according to another exemplary aspect of the present disclosure may include a core nacelle defined about an engine centerline axis, a fan nacelle mounted at least partially around the core nacelle to define a fan bypass flow path for a fan bypass airflow; a fan variable area nozzle axially movable relative the fan nacelle to define an auxiliary port to vary a fan nozzle exit area and adjust a pressure ratio of the fan bypass airflow during engine operation, and a controller operable to control the fan variable area nozzle to vary a fan nozzle exit area and adjust the pressure ratio of the fan bypass airflow. 
         [0025]    In a further non-limiting embodiment of any of the foregoing gas turbine embodiments, the gas turbine engine may be a direct drive turbofan engine. 
         [0026]    In a further non-limiting embodiment of any of the foregoing gas turbine embodiments, the gas turbine may further include a low spool within the core nacelle that drives a fan within the fan nacelle through a geared architecture. 
         [0027]    In a further non-limiting embodiment of any of the foregoing gas turbine embodiments, the engine may have a bypass ratio greater than 10:1 and the geared architecture may have a gear reduction ratio of greater than 2.5:1. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0028]    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 embodiment. The drawings that accompany the detailed description can be briefly described as follows: 
           [0029]      FIG. 1A  is a general schematic partial fragmentary view of an exemplary gas turbine engine embodiment for use with the present invention; 
           [0030]      FIG. 1B  is a rear view of the engine; 
           [0031]      FIG. 1C  is a side view of the engine integrated with a pylon; 
           [0032]      FIG. 1D  is a perspective view of the engine integrated with a pylon; 
           [0033]      FIG. 2A  is a sectional side view of the VAFN in a closed position; 
           [0034]      FIG. 2B  is a sectional side view of the VAFN in an open position; and 
           [0035]      FIG. 3  is a graph of a bypass duct normalized cross-sectional area distribution. 
           [0036]      FIG. 4  is a graph of a Effective Area Increase vs. Nozzle Translation; 
           [0037]      FIG. 5  is a graph of a duct area distribution; 
           [0038]      FIG. 6A  is schematic geometric view of the auxiliary port location; 
           [0039]      FIG. 6B  is schematic geometric view of the auxiliary port entrance angle; and 
           [0040]      FIG. 6C  is schematic geometric view of a VAFN outer surface curvature. 
       
    
    
     DETAILED DESCRIPTION 
       [0041]      FIG. 1A  illustrates a general partial fragmentary schematic view of a gas turbofan engine  10  suspended from an engine pylon P within an engine nacelle assembly N as is typical of an aircraft designed for subsonic operation. 
         [0042]    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  drives a fan section  20  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. 
         [0043]    The engine  10  is preferably a high-bypass geared architecture aircraft engine. In one disclosed, non-limiting embodiment, the engine  10  bypass ratio is greater than about six (6) to ten (10), the gear train  22  is an epicyclic gear train such as a planetary gear system or other gear system with a gear reduction ratio of greater than about 2.3 and the low pressure turbine  18  has a pressure ratio that is greater than about 5. In one disclosed embodiment, the engine  10  bypass ratio is greater than ten (10:1), the turbofan diameter is significantly larger than that of the low pressure compressor  16 , and the low pressure turbine  18  has a pressure ratio that is greater than 5:1. The gear train  22  may be an epicycle gear train such as a planetary gear system or other gear system with a gear reduction ratio of greater than 2.5:1. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans. 
         [0044]    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  to power the low pressure compressor  16  and the high pressure compressor  26 . Core airflow compressed by the low pressure compressor  16  and the high pressure compressor  26  is mixed with the fuel in the combustor  30  and expanded over the high pressure turbine  28  and low pressure turbine  18 . 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 . 
         [0045]    The core nacelle  12  is supported within the fan nacelle  34  by structure  36  often generically referred to as Fan Exit Guide Vanes (FEGVs). A bypass flow path  40  is defined between the core nacelle  12  and the fan nacelle  34 . The engine  10  generates a high bypass flow arrangement with a bypass ratio in which approximately 80 percent of the airflow entering the fan nacelle  34  becomes bypass flow B. The bypass flow B communicates through the generally annular fan bypass flow path  40  and is discharged from the engine  10  through a fan variable area nozzle (VAFN)  42  which defines a fan 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 . 
         [0046]    Thrust is a function of density, velocity, and area. One or more of these parameters can be manipulated to vary the amount and direction of thrust provided by the bypass flow B. The VAFN  42  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 pressure ratio turbofans are desirable for their high propulsive efficiency. However, low pressure ratio fans may be inherently 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 the instability region, and still provide the relatively smaller nozzle area necessary to obtain a high-efficiency fan operating line at cruise. 
         [0047]    A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  20  of the engine  10  is preferably designed for a particular flight condition—typically cruise at about 0.8M and about 35,000 feet. As the fan blades within the fan section  20  are efficiently designed at a particular fixed stagger angle for an efficient cruise condition, the VAFN  42  is 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 operational parameters such as noise levels. 
         [0048]    The VAFN  42  is separated into at least two sectors  42 A- 42 B ( FIG. 1B ) defined between the pylon P and a lower Bi-Fi splitter L which typically interconnects a larger diameter fan duct reverser cowl and a smaller diameter core cowl ( FIG. 1C and 1D ). Each of the at least two sectors  42 A- 42 B are independently adjustable to asymmetrically vary the fan nozzle exit area  44  to generate vectored thrust. It should be understood that although two segments are illustrated, any number of segments may alternatively or additionally be provided. 
         [0049]    In operation, the VAFN  42  communicates with a controller C or the like to adjust the fan nozzle exit area  44  in a symmetrical and asymmetrical manner. Other control systems including an engine controller or aircraft flight control system may also be usable with the present invention. By adjusting the entire periphery of the VAFN  42  symmetrically in which all sectors are moved uniformly, thrust efficiency and fuel economy are maximized during each flight condition. By separately adjusting the circumferential sectors  42 A- 42 B of the VAFN  42  to provide an asymmetrical fan nozzle exit area  44 , engine bypass flow is selectively vectored to provide, for example only, trim balance or thrust controlled maneuvering enhanced ground operations or short field performance. 
         [0050]    The VAFN  42  generally includes an auxiliary port assembly  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  slides aftward upon a track fairing  56 A,  56 B (illustrated schematically in  FIG. 1C and 1D ) in response to an actuator  58  (illustrated schematically). The track fairing  56 A,  56 B extend from the first fan nacelle section  52  adjacent the respective pylon P and the lower Bi-Fi splitter L ( FIG. 1D ). 
         [0051]    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. 2A ) and an open position ( FIG. 2B ). The auxiliary port assembly  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 F 0  ( FIG. 2A ). 
         [0052]    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 an auxiliary port  60  which extends between the open second fan nacelle section  54  relative the first fan nacelle section  52  to essentially provide an increased fan nozzle exit area  44  exit area F 1 . That is, the exit area F 1  with the port  60  is greater than exit area F 0  ( FIG. 2B ). 
         [0053]    In one disclosed embodiment, the auxiliary port  60  is incorporated into the exhaust system of a high bypass ratio commercial turbofan engine within the bypass duct aft of the Fan Exit Guide Vanes (FEGVs;  FIGS. 2A, 2B ). The auxiliary port  60  is located in the aft section of the bypass duct outer wall. 
         [0054]    Referring to  FIG. 3 , the bypass duct area distribution, the effective area increase vs. translation ( FIG. 4 ), area distribution ( FIG. 5 ), and auxiliary port  60  location ( FIG. 6A ) and wall curvatures ( FIG. 6B-6C ) are tailored to provide a proper flow-field that allows the auxiliary port  60  to obtain the required additional effective exit area. The auxiliary port  60  will essentially double the effective area gain due to translation. The auxiliary port  60  provides a relatively low weight method of providing increased exit area to control the fan operating line without causing high system losses or unacceptable aircraft installation issues. By tailoring the bypass duct area distribution and outer wall curvature, the desired maximum effective area increase is achieved before the stroke of the auxiliary port  60  reaches its effective area increase limit. 
         [0055]    The auxiliary port exit plane  44 B (defined as the plane between the stationary section&#39;s trailing edge and the moving sections leading edge) initially has an opening in which the exit plane normal vector is near-axial, but as the stroke increases, the normal vector becomes more inclined and approaches a near-radial vector. Once the exit plane normal has become near-radial, the maximum auxiliary port effectiveness has been reached. Once this point is reached, the rate of the effective area vs. translation changes from steep slope of the “well designed port” the shallow rate of the “main nozzle only”, since additional area will be provided through the main nozzle  44 A due to the inward slope of the core nacelle  12 . A well designed auxiliary port nozzle will achieve approximately +25% effective area before the port effectiveness limit is reached. That is, there is a limited range of stroke in which the auxiliary port doubles the rate of additional effectiveness. Outside of this range, the rate of additional effectiveness may be equivalent to a translating nozzle that has no auxiliary port. Or put another way, the auxiliary port reduces the stroke necessary for a pure translating nozzle to achieve a desired effective area. 
         [0056]    Referring to  FIG. 5 , the cross-sectional area at the auxiliary port  60  is greater than the maximum required effective area of the VAFN  42  and the bypass duct area distribution is tailored to ensure the duct cross-sectional area forward of the auxiliary port  60  is greater than the port opening cross-sectional area. This avoids a situation where an upstream internal cross-section becomes the controlling flow area (i.e. is smaller than the exit area), which can lead to operational limits and structural issues. 
         [0057]    Referring to  FIG. 6A , the auxiliary port  60  in the disclosed embodiment, is located no more forward than 0.1 DEL_X/L_DUCT defined from a point D at the largest radius Rmax of the annular fan bypass flow path  40  defined by the second fan nacelle section  54 . Rmax is defined through point D and perpendicular to the engine axis A. Point D in the disclosed non limiting embodiment is located on an inner wall surface  541  of the second fan nacelle section  54  when the second fan nacelle section  54  is in a closed position. DEL_X is the axial distance to the forward most point of the auxiliary port  60  from Rmax. L_DUCT is the overall axial length of the annular fan bypass flow path  40 . The angle between the mean port line and the fan duct outer wall is relatively low to provide well-behaved, low loss exit flow. In the disclosed embodiment, the auxiliary port  60  entrance angle (Theta_in) relative to the fan bypass duct OD wall, is less than 20 degrees ( FIG. 6B ) while the outer VAFN surface has an R_ARC/CHORD&gt;0.7 where R_ARC is a radial distance from the engine axis A to a radial outer wall surface  540  of the second fan nacelle section  54  and CHORD is the chord length of the second fan nacelle section  54  ( FIG. 6C ). The curvature of the outer wall surface  540  near the auxiliary port  60  promotes flow through the auxiliary port  60 . In one disclosed embodiment, the stroke of the second fan nacelle section  54  necessary to obtain an additional 20% effective exit area is approximately  8 . 4  inches. 
         [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 assembly  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 maximized during each flight regime by varying the fan nozzle exit area. By separately adjusting the sectors of the second fan nacelle section  54  to provide an asymmetrical fan nozzle exit area  44 , engine bypass flow is selectively vectored to provide, for example only, trim balance, thrust controlled maneuvering, enhanced ground operations and short field performance. 
         [0059]    The foregoing description is exemplary rather than defined by the limitations within. Many modifications and variations of the present invention are possible in light of the above teachings. The preferred embodiments of this invention 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. For that reason the following claims should be studied to determine the true scope and content of this invention.