Abstract:
A core nacelle for a gas turbine engine, according to an exemplary aspect of the present disclosure includes, among other things, a core cowl positioned adjacent to an inner duct boundary of a fan bypass passage having an associated cross-sectional area that radially extends between a fan exhaust nozzle and the inner duct boundary. The core cowl includes at least one groove that is selectively exposed to change the cross-sectional area at an axial location of the fan exhaust nozzle.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application is a continuation of U.S. patent application Ser. No. 12/444,487, filed on Apr. 6, 2009, which is the national stage entry of PCT/US06/39050, filed on Oct. 12, 2006. 
    
    
     BACKGROUND 
     This disclosure generally relates to a gas turbine engine, and more particularly to a turbofan gas turbine engine having a core nacelle including a corrugated core cowl. 
     In an aircraft gas turbine engine, such as a turbofan engine, air is pressurized in a compressor, and mixed with fuel and burned in a combustor for generating hot combustion gases. The hot combustion gases flow downstream through turbine stages that extract energy from the gases. A high pressure turbine powers the compressor, while a low pressure turbine powers a fan disposed upstream of the compressor. 
     Combustion gases are discharged from the turbofan engine through a core exhaust nozzle, and fan air is discharged through an annular fan exhaust nozzle defined at least partially by a fan nacelle surrounding the core engine. A significant amount of propulsion thrust is provided by the pressurized fan air which is discharged through the fan exhaust nozzle. The combustion gases are discharged through the core exhaust nozzle to provide additional thrust. 
     A significant amount of the air pressurized by the fan bypasses the engine for generating propulsion thrust in turbofan engines. High bypass turbofans typically require large diameter fans to achieve adequate turbofan engine efficiency. Therefore, the nacelle of the turbofan engine must be large enough to support the large diameter fan of the turbofan engine. Disadvantageously, the relatively large size of the nacelle results in increased weight, noise and drag that may offset the propulsive efficiency achieved by the high bypass turbofan engine. 
     It is known in the field of aircraft gas turbine engines that the performance of the turbofan engine varies during diverse flight conditions experienced by the aircraft. Typical turbofan engines are designed to achieve maximum performance during normal cruise operation of the aircraft. Therefore, when combined with the necessity of a relatively large nacelle size, increased noise and decreased efficiency may be experienced by the aircraft at non-cruise operability conditions such as take-off, landing, cruise maneuver and the like. 
     Accordingly, it is desirable to provide a turbofan engine having a variable discharge airflow cross-sectional area that achieves noise reductions and improved fuel economy in a relatively inexpensive and non-complex manner. 
     SUMMARY 
     A core nacelle for a gas turbine engine, according to an exemplary aspect of the present disclosure includes, among other things, a core cowl positioned adjacent to an inner duct boundary of a fan bypass passage having an associated cross-sectional area that radially extends between a fan exhaust nozzle and the inner duct boundary. The core cowl includes at least one groove that is selectively exposed to change the cross-sectional area at an axial location of the fan exhaust nozzle. 
     In a further non-limiting embodiment of the foregoing core nacelle, the at least one groove has an open position corresponding to a first cross-sectional area and a closed position corresponding to a second cross-sectional area less than the first cross-sectional area. 
     In a further non-limiting embodiment of either of the foregoing core nacelles, the at least one groove is sealed from a fan discharge airflow in the closed position and is exposed to the fan discharge airflow in the open position. 
     In a further non-limiting embodiment of any of the foregoing core nacelles, the at least one groove includes a plurality of grooves, the plurality of grooves each individually disposed circumferentially about an interior surface of the core cowl. 
     In a further non-limiting embodiment of any of the foregoing core nacelles, the core cowl includes an outer sleeve having a plurality of flap sections, wherein each of the plurality of flap sections are selectively moveable to expose the at least one groove. 
     In a further non-limiting embodiment of any of the foregoing core nacelles, the plurality of flap sections are stored within a cavity to expose the at least one groove. 
     In a further non-limiting embodiment of any of the foregoing core nacelles, the plurality of flap sections are circumferentially rotatable about an engine centerline axis. 
     In a further non-limiting embodiment of any of the foregoing core nacelles, the at least one groove includes a corrugation. The at least one groove is generally crescent shaped. 
     In a further non-limiting embodiment of any of the foregoing core nacelles, the at least one groove is formed on an interior surface of the core cowl at a section of the core cowl that is directly adjacent to an aftmost segment of the fan exhaust nozzle. 
     In a further non-limiting embodiment of any of the foregoing core nacelles, a radially inner portion of the at least one groove is generally crescent shaped. 
     In a further non-limiting embodiment of any of the foregoing core nacelles, the core cowl includes an outer sleeve and the at least one groove extends radially inwardly from the outer sleeve. 
     A gas turbine engine system, according to an exemplary aspect of the present disclosure includes, among other things, a fan nacelle defined about an axis and having a fan exhaust nozzle and a core nacelle at least partially within the fan nacelle. The core nacelle has a core cowl including at least one groove, the at least one groove defined on the core cowl at an axial location of the fan exhaust nozzle. The core cowl is selectively moveable between a first position having a first discharge airflow cross-sectional area and a second position having a second discharge airflow cross-sectional area greater than the first discharge airflow cross-sectional area. A fan section is positioned within the fan nacelle. At least one compressor and at least one turbine is positioned downstream of the fan section. At least one combustor is positioned between the at least one compressor and the at least one turbine. At least one sensor produces a signal representing an operability condition. A controller receives the signal. The controller selectively moves the core cowl from the first position to the second position in response to the signal. 
     In a further non-limiting embodiment of the foregoing gas turbine engine, the operability condition includes at least one of a take-off condition, an approach condition and a climb-condition. 
     In a further non-limiting embodiment of either of the foregoing gas turbine engines, the at least one groove is exposed to a fan discharge airflow in response to movement of the core cowl to the second position. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engines, an actuator assembly is in communication with the controller and operable to move the core cowl between the first position and the second position in response to the signal. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engines, the fan exhaust nozzle is positioned adjacent an aftmost segment of the fan nacelle. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engines, the at least one groove is formed on an interior surface of the core cowl at a section of the core cowl that is directly adjacent to the aftmost segment. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engines, a radially inner portion of the at least one groove is generally crescent shaped. 
     In a further non-limiting embodiment of any of the foregoing gas turbine engines, the core cowl includes an outer sleeve and the at least one groove extends radially inwardly from the outer sleeve. 
     The various features and advantages of this disclosure will become apparent to those skilled in the art from the following detailed description. The drawings that accompany the detailed description can be briefly described as follows. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  illustrates a general perspective view of an example gas turbine engine; 
         FIG. 2  illustrates an example core cowl in a closed position taken along section A-A of  FIG. 1 ; 
         FIG. 3  illustrates the example core cowl in an open position to expose a plurality of corrugations taken along section A-A of  FIG. 1 ; and 
         FIG. 4  shows an exploded view of a section of a slideable outer sleeve of the example core cowl for exposing the plurality of corrugations. 
     
    
    
     DETAILED DESCRIPTION 
     Referring to  FIG. 1 , a gas turbine engine  10  suspends from an engine pylon  12  as is typical of an aircraft designed for subsonic operation. In one example, the gas turbine engine is a geared turbofan aircraft engine. The gas turbine engine  10  includes a fan section  14 , a low pressure compressor  15 , a high pressure compressor  16 , a combustor  18 , a high pressure turbine  20  and a low pressure turbine  22 . A low speed shaft  19  rotationally supports the low pressure compressor  15  and the low pressure turbine  22  and drives the fan section  14  through a gear train  23 . A high speed shaft  21  rotationally supports the high pressure compressor  16  and a high pressure turbine  20 . The low speed shaft  19  and the high speed shaft  21  rotate about a longitudinal centerline axis A of the gas turbine engine  10 . 
     During operation, air is pressurized in the compressors  15 ,  16  and mixed with fuel and burned in the combustor  18  for generating hot combustion gases. The hot combustion gases flow through the high and low pressure turbines  20 ,  22  which extract energy from the hot combustion gases. 
     The example gas turbine engine  10  is in the form of a high bypass ratio (i.e., low fan pressure ratio geared) turbofan engine mounted within a fan nacelle  26 , in which most of the air pressurized by the fan section  14  bypasses the core engine itself for the generation of propulsion thrust. The example illustrated in  FIG. 1  depicts a high bypass flow arrangement in which approximately 80% of the airflow entering the fan nacelle  26  may bypass the core nacelle  28  via a fan bypass passage  27 . The high bypass flow arrangement provides a significant amount of thrust for powering the aircraft. 
     In one example, the bypass ratio is greater than ten, and the fan section  14  diameter is substantially larger than the diameter of the low pressure compressor  15 . The low pressure turbine  22  has a pressure ratio that is greater than five, in one example. The gear train  23  can be any known gear system, such as a planetary gear system with orbiting planet gears, planetary system with non-orbiting planet gears, or other type of gear system. In the disclosed example, the gear train  23  has a constant gear ratio. It should be understood, however, that the above parameters are only exemplary of a contemplated geared turbofan engine. That is, the invention is applicable to other engine architectures. 
     A fan discharge airflow F 1  is communicated within the fan bypass passage  27  and is discharged from the engine  10  through a fan exhaust nozzle  30 , defined radially between a core nacelle  28  and the fan nacelle  26 . Core exhaust gases C are discharged form the core nacelle  28  through a core exhaust nozzle  32  defined between the core nacelle  28  and a tail cone  34  disposed coaxially therein around the longitudinal centerline axis A of the gas turbine engine  10 . 
     The fan exhaust nozzle  30  concentrically surrounds the core nacelle  28  near an aftmost segment  29  of the fan nacelle  26 , in this example. In other examples, the fan exhaust nozzle  30  is located farther upstream but aft of the fan section  14 . The fan exhaust nozzle  30  defines a discharge airflow cross-sectional area  36  between the fan nacelle  26  and the core nacelle  28  for axially discharging the fan discharge airflow F 1  pressurized by the upstream fan section  14 . The core nacelle  28  of the gas turbine engine  10  includes a core cowl  38 . The core cowl  38  represents an exterior flow surface of a section of the core nacelle  28 . The core cowl  38  is positioned adjacent an inner duct boundary  25  of the fan bypass passage  27 . 
       FIG. 2  illustrates an example arrangement of the core cowl  38 . In this example, the core cowl  38  is in a closed position. An interior surface  40  of the core cowl  38  includes a plurality of grooves such as corrugations  42 , for example. In one example, the corrugations  42  are generally crescent shaped. Although the example core cowl  38  is shown and described as having corrugations, it should be understood that the core cowl  38  may be designed including any other fluid channeling features to effectively increase the discharge airflow cross-sectional area  36 . That is, the corrugations  42  provide additional area for the fan discharge airflow F 1  to flow over the core cowl  38 , as is further discussed below. 
     The plurality of corrugations  42  are manufactured from the same material as the core cowl  38 . The plurality of corrugations  42  are individually disposed circumferentially about the interior surface  40  of the core cowl  38 . That is, the plurality of corrugations  42  are not connected to one another. 
     In the closed position, the plurality of corrugations  42  are not exposed to the fan discharge airflow F 1 . Therefore, in this example, the discharge airflow cross-sectional area  36  extends between the aftmost segment  29  of the fan nacelle  26  and an outer sleeve  44  of the core cowl  38 , as is further discussed below. In one example, the plurality of corrugations  42  are formed on the interior surface  40  of the core cowl  38  at a section of the core cowl  38  that is directly adjacent to an aftmost segment of the fan exhaust nozzle  30 . However, the actual size, shape and location of the plurality of corrugations  42  will vary depending upon design specific parameters including, but not limited to, the size of the core nacelle  28  and the efficiency requirements of the gas turbine engine  10 . 
       FIG. 3  shows the core cowl  38  in an open (i.e., actuated) position. 
     Opening the core cowl  38  to expose the plurality of corrugations  42  during specific flight conditions provides noise reductions and improved fuel economy of the gas turbine engine  10 . In one example, the discharge airflow cross-sectional area  36  of the gas turbine engine  10  is varied by opening the core cowl  38  between the closed position ( FIG. 2 ) and the open position. The plurality of corrugations  42  are exposed to the fan discharge airflow F 1  by an actuator assembly  52  (See  FIG. 4 ) in response to detecting an operability condition. 
     In one example, the operability condition includes at least one of a take-off condition, an approach condition and a climb condition. Take-off conditions are experienced as the aircraft travels down the runway just prior to becoming airborne. Approach conditions are experienced during aircraft descent toward a landing strip to land the aircraft. Climb conditions are experienced where an aircraft reaches a certain altitude and cuts back against oncoming airflow to begin normal cruise operation. However, the plurality of corrugations  42  may be exposed in response to any known operability condition. 
     A discharge airflow cross-sectional area  46  associated with the opened core cowl  38  is greater than the discharge airflow cross-sectional area  36  of the core cowl  38  in its closed position. The discharge airflow cross-sectional area  46  includes the area defined by the discharge airflow cross-sectional area  36  and an area AR defined by each corrugation  42  to provide an increased airflow cross-sectional area for the fan discharge airflow F 1 . The actual size of the area AR of each corrugation  42  will depend upon design specific parameters including, but not limited to, the actual size and performance requirements of the gas turbine engine  10 . 
     A sensor  48  detects the operability condition and communicates a signal to a controller  50  to open the core cowl  38  and expose the plurality of corrugations  42  via an actuator assembly  52 . Of course, this view is highly schematic. It should be understood that the sensor  48  and the controller  50  may be programmed to detect known operability conditions. A person of ordinary skill in the art having the benefit of the teachings herein would be able to program the controller  50  to communicate with the actuator assembly  52  to move the core cowl  38  between the closed position and the open position. The actuator assembly  52  returns the core cowl  38  to the closed position, and the plurality of corrugations  42  are sealed from exposure to the fan discharge airflow F 1 , during normal cruise operation (e.g., a generally constant speed at generally constant, elevated altitude) of the aircraft. 
       FIG. 4  illustrates a section of the outer sleeve  44  of the core cowl  38 . The outer sleeve  44  is disposed coaxially about the core cowl  38  and includes a plurality of flap sections  54 . For simplification, only one flap section  54  is shown. The flap section  54  is selectively movable by the actuator assembly  52  to expose the corrugation  42  in response to detecting the operability condition. The flap section  54  is circumferentially rotatable about the engine centerline axis A. In one example, the flap section  54  slides in a clockwise direction. In another example, the flap section  54  moves in a counter-clockwise direction. In yet another example, the flap section  54  moves in both a clockwise and a counter-clockwise direction. 
     Each flap section  54  is stored within a cavity  60  of the outer sleeve  44  where the core cowl  38  is actuated to an open position. Therefore, the corrugations  42  are exposed to the fan discharge airflow F 1  and an increased discharge airflow cross-sectional area is achieved. The increase in the discharge airflow cross-sectional area enables noise reductions and improves fuel economy of the gas turbine engine  10 . In addition, control of the discharge airflow cross-sectional area provides control of the pressure of the gas turbine engine  10  within the fan bypass passage  27 , which in turn provides control over the fan pressure ratio of the gas turbine engine. The actuator assembly  52  moves the flap section  54  within the cavity  60  in response to detecting an operability condition. 
     The actuator assembly  52  extends each flap section  54  between adjacent sections of the outer sleeve  44  where an increase in the discharge airflow cross-sectional area is no longer desired (i.e., during normal cruise operation). 
     One example actuator assembly  52  is an electric actuation device. In another example, the actuator assembly  52  is a hydraulic actuation device. A worker of ordinary skill in the art with the benefit of the teachings herein would understand how to translate the flap sections  54  of the outer sleeve  44  to expose the plurality of corrugations  42  and provide an increased flow area for the fan discharge airflow F 1 . 
     The foregoing description shall be interpreted as illustrative and not in any limiting sense. A worker of ordinary skill in the art would recognize that certain modifications would come within the scope of this invention. For that reason, the following claims should be studied to determine the true scope and content of this invention.