Abstract:
A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a nacelle assembly that includes an inlet lip section and an inlet internal diffuser section downstream of the inlet lip section. A variable area fan nozzle is positioned near an aft segment of the nacelle assembly, the variable area fan nozzle adaptable to move between a first position having a first discharge airflow area and a second position having a second discharge airflow area greater than the first discharge airflow area. At least one boundary layer control device is positioned near one of the inlet lip section and the inlet internal diffuser section. A controller is configured to move the variable area fan nozzle from the first position to the second position and to actuate the at least one boundary layer control device to introduce an airflow in response to an operability condition.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application is a continuation of U.S. patent application Ser. No. 12/832,280, filed on Jul. 8, 2010, which is a divisional of U.S. patent application Ser. No. 11/584,030, which was filed on Oct. 20, 2006 and issued on Sep. 21, 2010 as U.S. Pat. No. 7,797,944. 
     
    
     BACKGROUND 
       [0002]    This disclosure generally relates to a gas turbine engine, and more particularly to a nacelle for a turbofan gas turbine engine. 
         [0003]    In an aircraft gas turbine engine, such as a turbofan engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. The hot combustion gases flow downstream through turbine stages that extract energy from the gases. In a two spool gas turbine engine, a high pressure turbine powers the high pressure compressor, while a low pressure turbine powers a fan disposed upstream of the compressor and a low pressure compressor. 
         [0004]    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 nacelle surrounding the core engine. A majority of propulsion thrust is provided by the pressurized fan air which is discharged through the fan exhaust nozzle, while the remaining thrust provided from the combustion gases is discharged through the core exhaust nozzle. 
         [0005]    In high bypass turbofans a majority of the air pressurized by the fan bypasses the turbofan engine for generating propulsion thrust. High bypass turbofans typically use 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. The relatively large size of the nacelle results in increased weight and drag that may offset the propulsive efficiency achieved by high bypass turbofan engines. 
       SUMMARY 
       [0006]    A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a nacelle assembly that includes an inlet lip section and an inlet internal diffuser section downstream of the inlet lip section. A variable area fan nozzle is positioned near an aft segment of the nacelle assembly, the variable area fan nozzle adaptable to move between a first position having a first discharge airflow area and a second position having a second discharge airflow area greater than the first discharge airflow area. At least one boundary layer control device is positioned near one of the inlet lip section and the inlet internal diffuser section. A controller is configured to move the variable area fan nozzle from the first position to the second position and to actuate the at least one boundary layer control device to introduce an airflow in response to an operability condition. 
         [0007]    In a further non-limiting embodiment of the foregoing gas turbine engine, the inlet internal diffuser section extends between a throat of an outer surface of the nacelle assembly and a forward face of a fan. 
         [0008]    In a further non-limiting embodiment of either of the foregoing gas turbine engines, the fan includes a variable pitch fan blade. 
         [0009]    In a further non-limiting embodiment of any of the foregoing gas turbine engines, the controller is configured to move the variable area fan nozzle and actuate the at least one boundary layer control device in response to the operability condition. 
         [0010]    In a further non-limiting embodiment of any of the foregoing gas turbine engines, the operability condition includes a windmilling condition. 
         [0011]    In a further non-limiting embodiment of any of the foregoing gas turbine engines, the second position of the variable area fan nozzle is at least 20% of the opening capability of the variable area fan nozzle. 
         [0012]    In a further non-limiting embodiment of any of the foregoing gas turbine engines, the second position of the variable area fan nozzle is at least 10% of the opening capability of the variable area fan nozzle. 
         [0013]    In a further non-limiting embodiment of any of the foregoing gas turbine engines, the at least one boundary layer control device includes a first boundary layer control device at the inlet lip section and a second boundary layer control device at the inlet internal diffuser section. 
         [0014]    In a further non-limiting embodiment of any of the foregoing gas turbine engines, the variable area fan nozzle includes a synchronizing ring, a static ring and a flap assembly. 
         [0015]    In a further non-limiting embodiment of any of the foregoing gas turbine engines, a sensor is configured to detect the operability condition. 
         [0016]    A gas turbine engine method according to another exemplary aspect of the present disclosure includes, among other things, sensing an operability condition, increasing a discharge airflow area of a variable area fan nozzle and introducing an airflow at a surface of a nacelle assembly. The steps of increasing and introducing are performed in response to the operability condition from the sensing step. 
         [0017]    In a further non-limiting embodiment of the foregoing method, the operability condition includes a windmilling condition. 
         [0018]    In a further non-limiting embodiment of either of the foregoing methods, the steps of increasing and introducing are performed in response to the operability condition from the sensing step. 
         [0019]    In a further non-limiting embodiment of any of the foregoing methods, the method includes returning the variable area fan nozzle to its original position during a second operability condition. 
         [0020]    In a further non-limiting embodiment of any of the foregoing methods, the method of increasing includes moving the variable area fan nozzle between a first position and a second position that is at least 20% of its opening capability if the operability condition is a windmilling condition or a static condition. 
         [0021]    In a further non-limiting embodiment of any of the foregoing methods, the method of increasing includes moving the variable area fan nozzle between a first position and a second position that is at least 10% of its opening capability if the operability condition is a crosswind condition or a high angle of attack condition. 
         [0022]    In a further non-limiting embodiment of any of the foregoing methods, the method of introducing is not performed if the operability condition is a cruise condition. 
         [0023]    A method of designing a gas turbine engine according to another exemplary aspect of the present disclosure includes, among other things, providing a nacelle assembly that includes an inlet lip section and an inlet internal diffuser section downstream of the inlet lip section, positioning a variable area fan nozzle near an aft segment of the nacelle assembly, the variable area fan nozzle adaptable to move between a first position having a first discharge airflow area and a second position having a second discharge airflow area greater than the first discharge airflow area. The method additionally includes positioning at least one boundary layer control device near one of the inlet lip section and the inlet internal diffuser section and moving the variable area fan nozzle from the first position to the second position and actuating the at least one boundary layer control device to introduce an airflow in response to an operability condition. 
         [0024]    In a further non-limiting embodiment of the foregoing method, the method includes sensing the operability condition prior to the moving step. 
         [0025]    In a further non-limiting embodiment of either of the foregoing methods, the method includes providing a controller configured to move the variable area fan nozzle and to actuate the at least one boundary layer control device. 
         [0026]    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 
         [0027]      FIG. 1  illustrates a general perspective view of a gas turbine engine; 
           [0028]      FIG. 2  is a schematic view of a gas turbine engine having a variable area fan nozzle (VAFN); 
           [0029]      FIG. 3  is a perspective view of a section of the VAFN; 
           [0030]      FIG. 4  illustrates a gas turbine engine having a VAFN and a first boundary layer control device; and 
           [0031]      FIG. 5  illustrates a gas turbine engine having a VAFN, a first boundary layer control device and a second boundary layer control device for achieving a slim line nacelle. 
       
    
    
     DETAILED DESCRIPTION 
       [0032]    Referring to  FIG. 1 , a gas turbine engine  10  typically includes (in serial flow communication) a fan  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 . During operation, air is pressurized in the compressors  15 ,  16  and mixed with fuel 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 high pressure turbine  20  powers the high pressure compressor  16  through a shaft defined therebetween, and the low pressure turbine  22  powers the fan  14  and the low pressure compressor  15  through another shaft defined therebetween. The disclosure is not limited to the two spool axial gas turbine architecture described and may be used with other architectures, such as a single spool axial design, a three spool axial design and other architectures. 
         [0033]    The gas turbine engine  10  is in the form of a high bypass ratio turbofan engine mounted within a nacelle assembly  26 , in which most of the air pressurized by the fan  14  bypasses the core engine itself for the generation of propulsion thrust. The nacelle assembly  26  includes a fan cowl  46  and a core cowl  28  within the fan cowl  46 . Fan discharge airflow F 1  is discharged from the engine  10  through a variable area fan nozzle (VAFN)  30  defined radially between the core cowl  28  and the fan cowl  46 . Core exhaust gases C are discharged from the core engine through a core exhaust nozzle  32  defined between the core cowl  28  and a center plug  34  disposed coaxially therein around a longitudinal centerline axis A of the gas turbine engine  10 . 
         [0034]    The VAFN  30  concentrically surrounds the core cowl  28  near an aftmost segment  29  of the nacelle assembly  26 . The VAFN  30  of the nacelle assembly  26  defines a fan-nozzle discharge airflow area  36  ( FIG. 2 ) between the fan cowl  46  and the core cowl  28  for axially discharging the fan discharge airflow F 1  pressurized by the upstream fan  14 . 
         [0035]    Referring to  FIG. 2 , the nacelle assembly  26  defines an inlet lip section  38  and an inlet internal diffuser section  40 . The inlet lip section  38  is positioned near a forward segment  31  of the fan cowl  46 . The inlet internal diffuser section  40  is defined between a throat  42  of the fan cowl  46  and a forward face of the fan  14 . The fan cowl  46  defines an outer surface of the nacelle assembly  26 . The nacelle assembly  26  also defines a highlight diameter D h  and a maximum diameter D max . The highlight diameter D h  represents the diameter defined by the inlet lip section  38  of the nacelle assembly  26 . The maximum diameter D. represents the peak diameter of the nacelle assembly  26 . The throat  42  of the nacelle assembly  26  also defines a throat diameter D t . 
         [0036]    The maximum diameter D. of the nacelle assembly  26  may be established by Extended-Range Twin-Engine Operational Performance Standards (ETOPS) requirements, in which an external airflow F 2  over the fan cowl  46  is required to remain separation free under an engine-out windmilling condition or other condition. ETOPS requirements are aircraft performance standards established by the International Civil Aviation Organization. It is desirable from an engine efficiency standpoint for the external airflow F 2  to maintain attached to the fan cowl  46  during aircraft operation. A windmilling condition occurs where an engine of a twin-engine aircraft loses functionality (i.e. engine out condition). The damaged engine is advantageously permitted to rotate, and is driven by an airflow resulting from the forward velocity of the aircraft (i.e., the damaged engine is permitted to “windmill”). 
         [0037]    A diameter ratio, or the ratio of the highlight diameter D h  to the maximum diameter D max , is utilized to determine whether the nacelle assembly  26  achieves this ETOPS requirement and maintains an external airflow F 2  which is separation free from the fan cowl  46 . Current industry standards typically use a diameter ratio of at least approximately 0.80 to achieve a separation free airflow, but other diameter ratios may be feasible. 
         [0038]    The nacelle assembly  26  also defines a contraction ratio. The contraction ratio represents a relative thickness of the inlet lip section  38  of the nacelle assembly  26  and is represented by the ratio of a highlight area H a  (ring-shaped area defined by highlight diameter D h ) and a throat area T a  (ring-shaped area defined by throat diameter D t ) of the nacelle assembly  26 . Current industry standards typically use a contraction ratio of approximately 1.300 to prevent the separation of the fan discharge airflow F 1  from an interior wall  59  of the fan cowl  46 , but other contraction ratios may be feasible. “Thick” inlet lip section designs, which are associated with large contraction ratios, increase the maximum diameter and increase the weight and the drag penalties associated with the nacelle assembly  26 . The nacelle assembly  26  further defines an inlet lip length L lip  and a fan duct length L fan . 
         [0039]    Increasing the fan discharge airflow F 1  during specific flight conditions allows the external airflow F 2  to remain separation free from the fan cowl  46  while achieving a slim-line nacelle design. In one example, the increased fan discharge airflow F 1  is achieved by providing the gas turbine engine  10  with a VAFN  30  and increasing the discharge airflow area  36  of the VAFN  30  during the specific flight conditions. 
         [0040]    In one example, the increase in the discharge airflow area  36  is achieved by opening the VAFN  30 . For example, the VAFN  30  generally includes a synchronizing ring  41 , a static ring  43 , and a flap assembly  45  (See  FIG. 3 ). Other VAFN designs and actuation mechanisms may be used. The flap assembly  45  is pivotally mounted to the static ring  43  at a multitude of hinges  47  and linked to the synchronizing ring  41  through a linkage  49 . An actuator assembly  51  (only one shown in  FIG. 3 ) selectively rotates the synchronizing ring  41  relative the static ring  43  to adjust the flap assembly  45  through the linkage  49 . The radial movement of the synchronizing ring  41  is converted to tangential movement of the flap assembly  45  to vary the discharge airflow area  36  defined by the VAFN  30  through which the fan discharge airflow F 1  is discharged. 
         [0041]    The increase in the discharge airflow area  36  is achieved by moving the VAFN  30  from a first position to a second (or open) position X (represented by dashed lines in  FIG. 2 ) in response to a detected windmilling condition. The discharge airflow area  36  of the second position is greater than the discharge airflow area  36  of the first position. A sensor  53  detects the windmilling condition and communicates with a controller  55  to move the VAFN  30  via the actuator assembly  51 . It should be understood that the sensor  53  and the controller  55  may be programmed to detect any known flight condition. In one example, the second position X represents moving the VAFN  30  to approximately 20% of its opening capability during the windmilling condition, although the actual percentage the VAFN  30  is opened will depend on design specific parameters of the gas turbine engine. A person of ordinary skill in the art would know how to design appropriate actuation and control systems to achieve comparable results with an alternative VAFN design. In another example, the increased fan discharge airflow F 1  is achieved by providing the gas turbine engine with a variable pitch fan blade. 
         [0042]    The opening of the VAFN  30  during windmilling conditions allows for a reduction in the maximum diameter D. of the nacelle assembly  26  while maintaining an external airflow F 2  which is separation free from the fan cowl  46 . Therefore, the nacelle assembly  26  achieves an improved (i.e. larger) diameter ratio. Further, the improved diameter ratio results in a weight savings and a reduction in nacelle drag (i.e., slim-line nacelle). The VAFN  30  is returned to its first position (represented by solid lines) during normal cruise operation of the aircraft. 
         [0043]    Referring to  FIG. 4 , a slim-line nacelle  50  is illustrated which includes a first boundary layer control device  52  in addition to the VAFN  30 . The slim line nacelle  50  offers additional nacelle drag and weight benefits over the nacelle assembly  26 . The first boundary layer control device  52  is positioned at the inlet lip section  38  of the slim line nacelle  50 . The first boundary layer control device  52  introduces an airflow F 4  near the inlet lip section  38  in a direction defined by an intake airflow F 3  prior to the onset of separation of the fan discharge airflow F 1  from the interior wall  59  of the slim line nacelle  50 . The first boundary layer control device  52  addresses any distortion associated with the fan discharge airflow F 1  as the fan discharge airflow F 1  is communicated from an upstream end of the engine  10  toward the downstream end. 
         [0044]    The first boundary layer control device  52  may introduce the airflow F 4  by injection or suction of airflow near the inlet lip section  38 . For example, fluid injection jet devices (for injection of airflow) or blowing slots (for suction of airflow) may be provided near the inlet lip section  38  to introduce the airflow F 4 . It should be understood that the nacelle may include any known boundary layer control technology. 
         [0045]    The first boundary layer control device  52  is actuated to generate the airflow F 4  in response to detection of at least one operability condition. The operability condition is detected by the sensor  53 . The sensor  53  communicates the detection of the operability condition to the controller  55 , which then actuates the first boundary layer control device  52  to generate the airflow F 4 . A person of ordinary skill in the art would understand how to program the sensor  53  and the controller  55  for performing these functions. 
         [0046]    In one example, the operability condition includes a static condition. Static conditions occur at low speeds (i.e., just prior to take-off). In another example, the operability condition includes a cross-wind condition. Cross-wind conditions are experienced during takeoff as the aircraft travels down the runway (i.e., where the aircraft experiences airflow in a roughly perpendicular direction with respect to the movement of the aircraft down the runway). In yet another example, the operability condition includes a high angle of attack condition. High angle of attack conditions are experienced where the aircraft is traveling at low speeds and the angle of incidence of the airflow relative to the inlet lip section  38  of the slim line nacelle  50  is relatively large. It should be understood that first boundary layer control device  52  may be controlled during any operability condition experienced by an aircraft during operation. 
         [0047]    In addition, the discharge airflow area  36  of the VAFN  30  may be increased simultaneously with the generation of the airflow F 4  by the first boundary layer control device  52  during the operability conditions to achieve further weight and drag reductions. In one example, both the VAFN  30  and the boundary layer control device  52  are utilized during all static conditions, cross-wind conditions, and high angle of attack conditions. The controller  55  is programmable to move the VAFN  30  to a position representing approximately 10% of its opening capability during cross-wind conditions and high angle of attack conditions, and to approximately 20% of its opening capability during static conditions. The first boundary layer control device  52  is turned off during windmilling conditions and during normal cruise operation of the aircraft to achieve optimal performance. 
         [0048]    The first boundary layer control device  52  and the VAFN  30  may be utilized simultaneously during the operability conditions to achieve a nacelle having a reduced contraction ratio while maintaining non-separation of the fan discharge airflow F 1  from the interior wall  59  of the slim line nacelle  50 . Therefore, corresponding weight and drag benefits are achieved by the slim-line nacelle  50 . 
         [0049]    Referring to  FIG. 5 , a second slim-line nacelle  58  is illustrated. The nacelle  58  includes a second boundary layer control device  60  in addition to the first boundary layer control device  52  and the VAFN  30 . The second boundary layer control device  60  is identical to the configuration of the first boundary layer control device  52  except that the second boundary layer control device  60  is positioned downstream from the first boundary layer control device and near the inlet internal diffuser section  40 . The second boundary layer control device  60  generates an airflow F 5  at the inlet internal diffuser section  40  to prevent separation of the fan discharge airflow F 1  from the interior wall  59  near this area of the nacelle  58 . 
         [0050]    The second boundary layer control device  60  is actuated by the controller  55  in response to detection of at least one operability condition. In one example, the second boundary layer control device  60  is utilized to generate the airflow F 5  during static conditions, cross-wind conditions, and high angle of attack conditions. Utilization of the second boundary layer control device  60  at the inlet internal diffuser section  40  of the nacelle  58  enables a reduction in the inlet lip length L lip  and the fan duct length L fan , thereby enabling a weight reduction in the nacelle design. The second boundary layer control device  60  is shut off during windmilling conditions and during normal cruise operation of the aircraft. 
         [0051]    In one example, the VAFN  30 , the first boundary layer control device  52 , and the second boundary layer control device  60  are exploited simultaneously during at least one of the operability conditions. In another example, the VAFN  30 , the first boundary layer control device  52  and the second boundary layer control device  60  are simultaneously utilized during all static conditions, cross-wind conditions and high angle of attack conditions which are detected by the sensor  53 . The slim-line nacelle  58  achieves further drag reduction benefits in response to the simultaneous utilization of all three technologies during diverse flight requirements. 
         [0052]    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 disclosure. For that reason, the following claims should be studied to determine the true scope and content of this disclosure.