Patent Publication Number: US-11391240-B2

Title: Gas turbine engine bifurcation located fan variable area nozzle

Description:
REFERENCE TO RELATED APPLICATIONS 
     This application is a continuation of U.S. patent application Ser. No. 15/889,304, filed Feb. 6, 2018, which a continuation of U.S. patent application Ser. No. 13/343,964, filed Jan. 5, 2012, which is a continuation in part of U.S. patent application Ser. No. 12/441,546, filed Mar. 17, 2009. 
    
    
     BACKGROUND OF THE INVENTION 
     The present invention relates to a gas turbine engine, and more particularly to a turbofan engine having a bifurcation which effectively varies a fan nozzle exit area by adjusting a variable area flow system within the bifurcation to selectively vary the bypass area through which bypass flow may pass. 
     Conventional gas turbine engines 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 in series along a longitudinal axis and are enclosed in a nacelle. An annular stream of primary airflow passes through a radially inner portion of the fan section and through the core engine to generate primary thrust. 
     Combustion gases are discharged from the core engine through a primary airflow path and are exhausted through a core exhaust nozzle. An annular fan flow path, disposed radially outwardly of the primary airflow path, passes through a radial outer portion between a fan nacelle and a core nacelle and is discharged through an annular fan exhaust nozzle defined at least partially by the fan nacelle and the core nacelle to generate fan thrust. A majority of propulsion thrust is provided by the pressurized fan air discharged through the fan exhaust nozzle, the remaining thrust provided from the combustion gases discharged through the core exhaust nozzle. 
     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 typically associated with the use of a fan variable area nozzle may be negated. 
     SUMMARY OF THE INVENTION 
     A gas turbine engine according to an exemplary aspect of the present disclosure may include a core engine defined about an axis, a gear system driven by the core engine, the gear system defines a gear reduction ratio of greater than or equal to about 2.3, a fan driven by the gear system about the axis to generate a bypass flow, and a variable area flow system which operates to effect the bypass flow. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the variable area flow system may include an annular fan variable area nozzle (FVAN). 
     In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the gas turbine engine may include a gear system driven by the core engine to drive the fan. The gear system may define a gear reduction ratio of greater than or equal to about 2.5. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the gas turbine engine may include a gear system driven by the core engine to drive the fan. The gear system may define a gear reduction ratio of greater than or equal to 2.5. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the core engine may include a low pressure turbine which defines a pressure ratio that is greater than about five (5). 
     In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the core engine may include a low pressure turbine which defines a pressure ratio that is greater than five (5). 
     In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the bypass flow may define a bypass ratio greater than about six (6). 
     In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the bypass flow may define a bypass ratio greater than about ten (10). 
     In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the bypass flow may define a bypass ratio greater than ten (10). 
     In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the variable area flow system may operate to change a pressure ratio of the bypass flow. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the variable area flow system may operate to vary an area of a fan nozzle exit area for the bypass flow. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the fan may be defined for a predefined flight condition. Additionally or alternatively, the predefined flight condition may be about 0.8 MACH and about 35,000 feet. Additionally or alternatively, the predefined flight condition may be 0.8 MACH and 35,000 feet. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the fan may include fan blades designed at a particular fixed stagger angle related to the flight condition. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the variable area flow system may operate to adjust the bypass flow such that an angle of attack of the fan blades are maintained close to a design incidence at flight conditions other than the predefined flight condition. 
     A gas turbine engine according to another exemplary aspect of the present disclosure may include a core engine defined about an axis. The core engine may include a low pressure turbine which defines a pressure ratio that is greater than about five (5), a fan driven by the core engine about the axis to generate a bypass flow, and a variable area flow system which operates to effect the bypass flow. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the variable area flow system may include an annular fan variable area nozzle (FVAN). 
     In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the core engine may include a low pressure turbine which defines a pressure ratio that is greater than five (5). 
     In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the gas turbine engine may include a gear system driven by the core engine to drive the fan. The gear system may define a gear reduction ratio of greater than or equal to about 2.5. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the bypass flow may define a bypass ratio greater than about six (6). Additionally or alternatively, the bypass flow may define a bypass ratio greater than about ten (10). Additionally or alternative, the bypass flow may define a bypass ratio greater than ten (10). 
     In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the variable area flow system may operate to change a pressure ratio of the bypass flow. Additionally or alternatively, the variable area flow system may operate to vary an area of a fan nozzle exit area for the bypass flow. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the fan may be defined for a predefined flight condition. Additionally or alternatively, the flight condition may be about 0.8 MACH and about 35,000 feet. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the fan may include fan blades designed at a particular fixed stagger angle related to the predefined flight condition. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the variable area flow system may operate to adjust the bypass flow such that an angle of attack of the fan blades are maintained close to a design incidence at flight conditions other than the predefined flight condition. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       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: 
         FIG. 1  is a general schematic partial fragmentary view of an exemplary gas turbine engine embodiment for use with the present invention; and 
         FIG. 2  is a sectional view through an engine pylon of the engine of  FIG. 1  at line  2 - 2  to illustrate a variable area flow system. 
     
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT 
       FIG. 1  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. 
     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  connected to the low spool  14  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. 
     The engine  10  is preferably a high-bypass geared turbofan 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. Preferably, the engine  10  bypass ratio is greater than ten (10), the fan 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. The gear train  22  is preferably an epicyclic gear train such as a planetary gear system or other gear system with a gear reduction ratio of greater than 2.5. It should be understood, however, that the above parameters are only exemplary of various preferred geared turbofan engines and that the present invention is likewise applicable to other gas turbine engines. 
     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  where is ignited, and burned. The resultant high pressure combustor products are expanded through the high pressure turbine  28  and low pressure turbine  18 . The turbines  28 ,  18  are rotationally coupled to the compressors  26 ,  16  respectively to drive the compressors  26 ,  16  in response to the expansion of the combustor product. The low pressure turbine  18  also drives the fan section  20  through the gear train  22 . 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 . 
     The core nacelle  12  is supported within the fan nacelle  34  by a pylon structure often generically referred to as an upper bifurcation  36 U and lower bifurcation  36 L, however, other types of pylons and supports at various radial locations may likewise be usable with the present invention. 
     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 (circumferentially broken only by the bifurcations  36 U,  36 L) bypass flow path  40  and is discharged from the engine  10  through an annular fan variable area nozzle (FVAN)  42  which defines a variable fan nozzle exit area  44  between the fan nacelle  34  and the core nacelle  12 . The upper bifurcation  36 U and the lower bifurcation  36 L, although aerodynamically optimized (best seen in  FIG. 2 ), occupies some portion of the volume between the core nacelle  12  and the fan nacelle  34 . 
     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 upper bifurcation  36 U preferably includes a pylon variable area flow system  50  having a passage  56  defined between a pylon intake  52  and a pylon exhaust  54  to selectively vary the FVAN  42  area through which bypass flow B may pass. Preferably, both the pylon intake  52  and the pylon exhaust  54  are variable and controlled in response to a controller  58 . It should be understood that although the upper bifurcation  36 U is illustrated in the disclosed embodiment as having the pylon variable area flow passage  50 , the lower bifurcation as well as other pylon structures may likewise include such variable area flow systems. 
     Referring to  FIG. 2 , the pylon variable area flow system  50  changes the pressure ratio of the bypass flow B. That is, the nozzle exit area  44  is effectively varied in area by opening and closing the additional flow area of the pylon variable area flow system  50  to vary the bypass flow B. It should be understood that various actuators  64 ,  66  in communication with the controller  58  may be utilized to operate the pylon intake  52  and the pylon exhaust  54  in response to predetermined flight conditions. It should be understood that either of the pylon intake  52  and the pylon exhaust  54  may be fixed but it is preferred that both are adjustable in response to the controller  58  to control the flow area through the flow passage  56 . 
     The flow passage  56  is defined around a component duct  55  within the upper bifurcation  36 U which provides a communication path for wiring harnesses, fluid flow conduits and other components to the core nacelle  12  from, for example, the aircraft wing. It should be understood that various flow passage  56  paths will likewise be usable with the present invention. 
     The pylon intake  52  preferably includes an adjustable intake such as a louver system  60  with empirically-designed turning vanes which most preferably have a variation of height to minimize the “shadowing” effect created by each upstream louver relative the next downstream louver. 
     The pylon exhaust  54  preferably includes a variable nozzle  59 . The variable nozzle  59  may include doors, flaps, sleeves or other movable structure which control the volume of additional fan bypass flow B+ through the FVAN  42 . 
     The pylon variable area flow system  50  changes the physical area through which the bypass flow B may pass. 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 0.8 MACH and about 35,000 feet. It should be understood that other arrangements as well as essentially infinite intermediate positions are likewise usable with the present invention. 
     In operation, the pylon variable area flow system  50  communicates with the controller  58  to effectively vary the area of the fan nozzle exit area  44  through independent or coordinated operation of the pylon intake  52  and the pylon exhaust  54 . Other control systems including an engine controller, a flight control computer or the like may also be usable with the present invention. As the fan blades of fan section  20  are efficiently designed at a particular fixed stagger angle for the cruise condition, the pylon variable area flow system  50  is operated to vary the area of the fan nozzle exit area  44  to adjust fan bypass air flow such that the angle of attack or incidence of the fan blades are maintained close to the design incidence at other flight conditions, such as landing and takeoff as well as to meet other operational parameters such as noise level. Preferably, the pylon variable area flow system  50  is closed to define a nominal cruise position fan nozzle exit area  44  and is opened for other flight conditions. The pylon variable area flow system  50  preferably provides an approximately 20% (twenty percent) effective area change in the fan nozzle exit area  44 . 
     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.