Abstract:
A method of improving the aerodynamic performance of an inlet section of a nacelle assembly of a gas turbine engine, according to an exemplary aspect of the present disclosure includes, among other things, (a) sensing an operability condition, (b) adjusting a leading edge of the inlet section in response to the step (a), and (c) adjusting a thickness of the inlet section in response to the step (a).

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application is a divisional of U.S. patent application Ser. No. 11/769,749, which was filed on Jun. 28, 2007. 
     
    
     BACKGROUND 
       [0002]    This disclosure generally relates to a gas turbine engine, and more particularly to a gas turbine engine having a variable shape inlet section. 
         [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 which extract energy from the hot combustion gases. A fan section supplies air to the compressor. 
         [0004]    Combustion gases are discharged from the turbofan engine through a core exhaust nozzle and a quantity of fan air is discharged through an annular fan exhaust nozzle defined at least partially by a nacelle assembly 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 is provided by the combustion gases discharged through the core exhaust nozzle. 
         [0005]    It is known in the field of aircraft gas turbine engines that the performance of a turbofan engine varies during diversified operability conditions experienced by the aircraft. An inlet lip section located at the foremost end of the turbofan nacelle assembly is typically designed to enable operation of the turbofan engine and reduce separation of airflow from the internal and external flow surfaces of the inlet lip section during these diversified conditions. For example, the nacelle assembly requires a “thick” inlet lip section to support operation of the engine during specific flight conditions, such as crosswind conditions, take-off conditions and the like. Disadvantageously, the “thick” inlet lip section may reduce the efficiency of the turbofan engine during normal cruise conditions of the aircraft, for example. As a result, the maximum diameter of the nacelle assembly is approximately 10-20% larger than required during cruise conditions. Since aircraft typically operate in cruise conditions for extended periods, turbofan efficiency gains can lead to substantially reduced fuel burn and emissions. 
         [0006]    Accordingly, it is desirable to provide a nacelle assembly having an adaptive structure to improve the performance of a turbofan gas turbine engine during diversified operability conditions. 
       SUMMARY 
       [0007]    A method of improving the aerodynamic performance of an inlet section of a nacelle assembly of a gas turbine engine, according to an exemplary aspect of the present disclosure includes, among other things, sensing an operability condition, adjusting a leading edge of the inlet section in response to the step of sensing, and adjusting a thickness of the inlet section in response to the step of sensing. 
         [0008]    In a further non-limiting embodiment of the foregoing method, the leading edge is adjustable between a thin position and a blunt position and the step of adjusting the leading edge includes the step of moving the leading edge between the thin position and the blunt position. 
         [0009]    In a further non-limiting embodiment of either of the foregoing methods, the thickness of the inlet section is adjustable between a first position and a second position and the step of adjusting the thickness includes adjusting the thickness between the first position and the second position. 
         [0010]    In a further non-limiting embodiment of any of the foregoing methods, the second position is radially outward from the first position. 
         [0011]    In a further non-limiting embodiment of any of the foregoing methods, the second position is radially inward from the first position. 
         [0012]    In a further non-limiting embodiment of any of the foregoing methods, the inlet section includes a plurality of discrete sections circumferentially disposed about an engine longitudinal centerline axis and each having a leading edge and a body panel portion aft of the leading edge that includes an adaptive structure. The step of adjusting the thickness includes adjusting a thickness of each of the body panel portions in each of a radially outer direction and a radially inner direction relative to the engine longitudinal centerline axis to alter the adaptive structure. 
         [0013]    In a further non-limiting embodiment of any of the foregoing methods, the adaptive structure of a first discrete section of the plurality of discrete sections is altered independently of the adaptive structure of a second discrete section of the plurality of discrete sections. 
         [0014]    In a further non-limiting embodiment of any of the foregoing methods, the step of sensing is performed using a programmable controller. 
         [0015]    A method of influencing an adaptive structure of a plurality of discrete sections of a gas turbine engine according to another exemplary aspect of the present disclosure includes, among other things, positioning the plurality of discrete sections circumferentially about an engine longitudinal centerline axis. Each of the plurality of discrete sections have a leading edge and a body panel portion aft of the leading edge that includes the adaptive structure. The method includes adjusting a thickness of at least one body panel portion in each of a radially outer direction and a radially inner direction relative to the engine longitudinal centerline axis to alter the adaptive structure. 
         [0016]    In a further non-limiting embodiment of the foregoing method, the method includes the step of identifying an operability condition prior to the step of adjusting. 
         [0017]    In a further non-limiting embodiment of either of the foregoing methods, the method includes the step of moving the leading edge between a thin position and a blunt position. 
         [0018]    In a further non-limiting embodiment of any of the foregoing methods, the adaptive structure of a first discrete section of the plurality of discrete sections is altered independently of the adaptive structure of a second discrete section of the plurality of discrete sections. 
         [0019]    In a further non-limiting embodiment of any of the foregoing methods, the thickness of the first discrete section and the thickness of the second discrete section are adjusted uniformly. 
         [0020]    In a further non-limiting embodiment of any of the foregoing methods, the thickness of the first discrete section and the thickness of the second discrete section are adjusted by a different thickness amount. 
         [0021]    In a further non-limiting embodiment of any of the foregoing methods, the plurality of discrete sections include at least a first discrete section and a second discrete section, and the step of adjusting the thickness is performed on one of the first discrete section and the second discrete section but not on the other of the first discrete section and the second discrete section. 
         [0022]    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 
         [0023]      FIG. 1  illustrates a general sectional view of a gas turbine engine; 
           [0024]      FIG. 2  illustrates a nacelle assembly of a gas turbine engine illustrated in  FIG. 1 ; 
           [0025]      FIG. 3  illustrates a general perspective view of the nacelle assembly of a gas turbine engine shown in  FIG. 1 ; 
           [0026]      FIG. 4A  illustrates a first example position of a leading edge of an inlet section of the nacelle assembly; 
           [0027]      FIG. 4B  illustrates a second example position of the leading edge of the inlet section of the nacelle assembly; 
           [0028]      FIG. 5  illustrates an example mechanism for manipulating an adaptive structure of an inlet section of a nacelle assembly; and 
           [0029]      FIG. 6  illustrates a side view of the inlet section of the nacelle assembly of a gas turbine engine. 
       
    
    
     DETAILED DESCRIPTION 
       [0030]      FIG. 1  illustrates a gas turbine engine  10  which includes (in serial flow communication) 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 . During operation, air is pulled into the gas turbine engine  10  by the fan section  14 , pressurized by the compressors  15 ,  16  and is mixed with fuel and burned in a combustor  18 . Hot combustion gases generated within the combustor  18  flow through the high and low pressure turbines  20 ,  22  which extract energy from the hot combustion gases. 
         [0031]    In a two-spool gas turbine engine architecture, the high pressure turbine  20  utilizes the energy extracted from the hot combustion gases to power the high pressure compressor  16  through a high speed shaft  19 , and the low pressure turbine  22  utilizes the energy extracted from the hot combustion gases to power the low pressure compressor  15  and the fan section  14  though a low speed shaft  21 . However, the invention is not limited to the two-spool gas turbine engine architecture described and may be used with other architectures, such as a single-spool axial design, a three-spool axial design and other architectures. That is, the present invention is applicable to any gas turbine engine, and to any application. 
         [0032]    The example gas turbine engine  10  is in the form of a high bypass ratio turbofan engine mounted within a nacelle assembly  26 , in which a significant amount of air pressurized by the fan section  14  bypasses the core engine  39  for the generation of propulsion thrust. The nacelle assembly  26  partially surrounds an engine casing  31  that houses the core engine  39  and its components. The airflow entering the fan section  14  may bypass the core engine  39  via a fan bypass passage  30  which extends between the nacelle assembly  26  and the engine casing  31  for receiving and communicating a discharge airflow F 1 . The high bypass flow arrangement provides a significant amount of thrust for powering the aircraft. 
         [0033]    The engine  10  may include a geartrain  23  that controls the speed of the rotating fan section  14 . The geartrain  23  can be any known gear system, such as a planetary gear system with orbiting planet gears, a planetary gear system with non-orbiting planet gears or other type of gear system. In the disclosed example, the geartrain  23  has a constant gear ratio. It should be understood, however, that the above parameters are only examples of a contemplated geared turbofan engine  10 . That is, the invention is applicable to traditional turbofan engines as well as other engine architectures. 
         [0034]    The discharge airflow F 1  is discharged from the engine  10  through a fan exhaust nozzle  33 . Core exhaust gases C are discharged from the core engine  39  through a core exhaust nozzle  32  disposed between the engine casing  31  and a center plug  34  disposed coaxially around a longitudinal centerline axis A of the gas turbine engine  10 . 
         [0035]      FIG. 2  illustrates an example inlet lip section  38  of the nacelle assembly  26 . The inlet lip section  38  is positioned near a forward segment  29  of the nacelle assembly  26 . A boundary layer  35  is associated with inlet lip section  38 . The boundary layer  35  represents an area adjacent to each of an inner and outer flow surface of the inlet lip section  38  at which the velocity gradient of airflow is zero. That is, the velocity profile of oncoming airflow F 2  goes from a free stream away from the boundary layer  35  to near zero at the boundary layer  35  due to friction forces that occur as the oncoming airflow F 2  passes over the outer and inner flow surfaces of the inlet lip section  38 . 
         [0036]    The inlet lip section  38  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 r ). Currently industry standards typically require a contraction ratio of approximately 1.33 to reduce the separation of oncoming airflow F 2  from the outer and inner flow surfaces of the inlet lip section  38  during engine operation, but other contraction ratios may be feasible. “Thick” inlet lip section designs, which are associated with large contraction ratios, increase the maximum diameter D max  and increase the weight and drag penalties associated with the nacelle assembly  26 . In addition, a desired ratio of the maximum diameter Dmax relative to the highlight diameter D h  is typically less than or equal to about 1.5, for example. A person of ordinary skill in the art would understand that other ratios of the maximum diameter Dmax relative to the highlight diameter D h  are possible and will vary depending upon design specific parameters. 
         [0037]    Referring to  FIG. 3 , the inlet lip section  38  includes a plurality of discrete sections  40  disposed circumferentially about the engine longitudinal centerline axis A. Each of the discrete sections  40  includes a leading edge  42  and a body panel portion  44 . Each discrete section  40  has an adaptive structure that is capable of a shape change. The inlet lip section  38  is sectioned into the plurality of discrete sections  40  to reduce the stiffness of the closed annular structure of the inlet lip section  38  and allow flexure thereof. Each discrete section  40  is designed to be capable of deformation (i.e., the materials remain within their elastic limits), yet simultaneously have the requisite stiffness to maintain a deformed shape while under aerodynamic and external pressure loads. In addition, as would be understood by those of ordinary skill in the art having the benefit of this disclosure, each discrete section  40  could slightly overlap with adjacent discrete sections  40  to allow the shape change of the inlet lip section  38  to occur without interference. A fixed nacelle portion  41  is positioned downstream from the inlet lip section  38 . 
         [0038]    In one example, the discrete sections  40  are comprised of an aluminum alloy. In another example, the discrete sections are comprised of a titanium alloy. It should be understood that any deformable material may be utilized to form the discrete sections  40 . A person of ordinary skill in the art having the benefit of this description would be able to choose an appropriate material for the example discrete sections  40  of the inlet lip section  38 . 
         [0039]    Influencing the adaptive structure of the inlet lip section  38  during specific flight conditions to achieve a desired shape change increases the amount of airflow communicated through the gas turbine engine  10  and reduces the internal and external drag experienced by the inlet lip section  38 . In one example, the adaptive structure of the inlet lip section  38  is influenced by adjusting the shape of the leading edge  42  of each discrete section  40  (see  FIGS. 4A and 4B ). In another example, the adaptive structure of the inlet lip section  38  is influenced by adjusting a thickness of the body panel portions  44  of each discrete section  40  (see  FIG. 5 ). In yet another example, the adaptive structure of the inlet lip section  38  is influenced by adjusting both the leading edge  42  and the thickness of the body panel portion  44  of each discrete section  40 . 
         [0040]      FIGS. 4A and 4B  illustrate the adjustment of the leading edge  42  of a discrete section  40  of the inlet lip section  38  between a first position X (see  FIG. 4A ) and a second position X′ (see  FIG. 4B ). The first position X represents a “thin” inlet lip section  38 . The second position X′ represents a “blunt” inlet lip section  38 . Each leading edge  42  is moved between the first position X and the second position X′ via a rotary actuator  46 , for example. The rotary actuator  46  rotates in either a clockwise or counterclockwise direction to move a linkage assembly  48  and adjust the leading edge  42  between the first position X and the second position X′. The rotary actuator  46  and the linkage assembly  48  are mounted within a cavity  50  of each discrete section  40 . 
         [0041]    At least one linkage assembly  48  is provided within each discrete section  40  and includes a plurality of linkage arms  52  and a plurality of pivot points  54 . The rotary actuator  46  pivots, toggles, extends and/or flexes the linkage arms  52  of the linkage assembly  48  about the pivot points  54  to move the leading edge  42  between the “thin”, first position X and the “blunt”, second position X′. Although the present example is illustrated with a rotary actuator and linkage arms connected via pivot points, other mechanisms may be utilized to move the leading edges  42  of the discrete sections  40  between the first position X and the second position X′, including but not limited to linear actuators, bell cranks, etc. A person of ordinary skill in the art having the benefit of this disclosure will be able to implement an appropriate actuator assembly to manipulate the leading edge  42  of each discrete section  40 . In addition, it should be understood that the leading edge  42  is moveable to any position between the first position X and second position X′. 
         [0042]    The adaptive structure of the inlet lip section  38  is influenced by moving the leading edge  42  of each discrete section  40  between the first position X and the second position X′ in response to detecting an operability condition of the gas turbine engine  10 . In one example, the operability condition includes a take-off condition. In another example, the operability condition includes a climb condition. In yet another example, the operability condition includes a landing condition. In still another example, the operability condition includes a high angle of attack condition. It should be understood that the adaptive structure of the inlet lip section  38  is adjustable in response to any operability condition experienced by the aircraft. Each leading edge  42  is positioned at/returned to the first position X during normal cruise conditions of the aircraft. 
         [0043]    A sensor  61  detects the operability condition and communicates with a controller  62  to translate the leading edge  42  between the first position X and the second position X′ and influence the adaptive structure of the inlet lip section  38 . Of course, this view is highly schematic. In addition, the illustrations of the movement of the inlet lip section  38  are shown exaggerated to better illustrate the adaptive structure thereof. A person of ordinary skill in the art would understand the distances the leading edge  42  should be moved between the position X and the second position X′ in response to sensing a specific operability condition. 
         [0044]    It should be understood that the sensor  61  and the controller  62  may be programmed to detect any known operability condition and that each operability condition may be associated with a distinct position of the leading edge  42  of the inlet lip section  38 . That is, the sensor  61  and the controller  62  are operable to situate the leading edge  42  of each discrete section  40  at a position which corresponds to the operability condition that is detected. Also, the sensor can be replaced by any controller associated with the gas turbine engine  10  or an associated aircraft. In fact, the controller  62  itself can include the “sensor” and generate the signal to adjust the contour of the inlet lip section  38 . 
         [0045]      FIG. 5  illustrates the adjustment of a thickness T of a body panel portion  44  of each discrete section  40  to influence the adaptive structure of the inlet lip section  38 . The thickness T of the body panel  44  is adjustable between a “thin” inlet lip section  38  and a “thick” inlet lip section  38 , for example. An inner surface  70  and an outer surface  72  of each body panel portion  44  are moveable in a Y direction (i.e., radially outward) to adjust each discrete section  40  to a “thick” position. In addition, the inner and outer surfaces  70 ,  72  are moveable in a Z direction to adjust each discrete section to a “thin” position. 
         [0046]    The thickness T adjustment of each body panel portion  44  is achieved via a linear actuator  56  and a linkage assembly  58 . The linear actuator  56  and the linkage assembly  58  are received in the cavity  50  of each discrete section  40 . Although the present example is illustrated with a linear actuator and linkage arms connected via pivot points, other mechanisms may be utilized to adjust the thickness T of each body panel portion  44 . 
         [0047]    The linear actuator  56  includes an actuator arm  60  which is moveable in a R or L direction to move the linkage assembly  58  and thereby adjust the thickness of the body panel portion  44 . The linkage assembly  58  includes a plurality of linkages  64  and a plurality of pivot points  66 . The linear actuator  56  adjusts the thickness T of each body panel portion  44  by retracting, pivoting, toggling, extending and/or flexing the linkages  64  about each pivot point  66 . In one example, the actuator arm  60  of the linear actuator  56  moves in a R direction to retract the outer skin (i.e., move the outer skin in the Z direction) of the body panel portion  44  and provide a “thin” inlet lip section  38 . In another example, the actuator arm  60  of the linear actuator  56  is moved in a L direction to expand the outer skin (i.e., move the outer skin in the Y direction) of the body panel portion  44  and provide a “thick” inlet lip section  38 . That is, the thickness T of each body panel portion  44  is adjusted either radially outwardly or radially inwardly to provide a “thick” inlet lip section or a “thin” inlet lip section, respectively. 
         [0048]    The thickness of each discrete section  40  is adjusted in response to detecting an operability condition. In one example, the operability condition includes a take-off condition. In another example, the operating condition includes a climb condition. In another example, the operability condition includes a high angle of attack condition. In still another example, the operability condition includes a landing condition. It should be understood that the thickness of the body panel portion  44  may be adjusted to influence the adaptive structure of the inlet lip section  38  in response to any operability condition experienced by the aircraft. The thickness T is adjusted/returned to a “thin” position at cruise conditions of the aircraft. 
         [0049]    A sensor  61 , as is shown in  FIGS. 4   a  and  4   b , detects the operability condition and communicates with a controller  62  to adjust the thickness T of each discrete section  40 . Of course, this view is highly schematic. In addition, the illustrations of the movement of the inlet lip section  38  are shown exaggerated to better illustrate the adaptive structure thereof. A person of ordinary skill in the art would understand the distances the thickness T should be adjusted in response to sensing a specific operability condition. 
         [0050]    It should be understood that the sensor  61  and the controller  62  may be programmed to detect any known operability condition and that each operability condition may be associated with a distinct thickness T of the body panel portions  44  of the discrete sections  40 . That is, the sensor  61  and the controller  62  are operable to adjust the thickness T of each discrete section  40  to a position which corresponds to the operability condition that is detected. The thickness T of each discrete section  40  may be adjusted uniformly or differently about the circumference. In some instances, such as operating during strong cross-winds, for example, only certain discrete sections  40  may be adjusted, while other discrete sections  40  are left unchanged. Also, the sensor can be replaced by any controller associated with the gas turbine engine  10  or an associated aircraft. In fact, the controller  62  itself can generate the signal to adjust the contour of the inlet lip section  38 . 
         [0051]    Although illustrated in  FIGS. 4 and 5  as having only a single mechanism for adjusting the shape of the inlet lip section  38  (i.e., one of a rotary actuator  46  with a linkage assembly  48  or a linear actuator  56  with a linkage assembly  58 ), it should be understood that each discrete section  40  could include both types of mechanisms to achieve both a leading edge adjustment and a thickness adjustment of the inlet lip section  38 . A person of ordinary skill in the art having the benefit of this disclosure would be able to design the inlet lip section  38  to achieve a desired aerodynamic performance level. 
         [0052]    Influencing the adaptive structure of the inlet lip section  38  may also be achieved during diverse operating conditions by “drooping” a portion of the inlet lip section  38  relative to a remaining portion of the inlet lip section  38  (See  FIG. 6 ). In one example, a portion of the discrete sections  40  positioned near a top portion  80  of the inlet lip section  38  are translated in an X direction and a portion of discrete sections  40  positioned near a bottom portion  82  of the inlet lip section  38  are translated in a Y direction to create a droop angle D relative to a plane  84  defined by the foremost end  86  of the inlet lip section  38 . The translations of the discrete sections  40  in the X and Y directions are achieved via adjustment of linkage assembly  48  (See  FIGS. 4   a  and  4   b ), the linkage assembly  58  (See  FIG. 5 ) or a combination of both the linkage assembly  48  and the linkage assembly  58 . The droop angle D is between 2 to 6 degrees relative to the plane  84 , in one example. Although  FIG. 6  illustrates the “droop” of the bottom portion  82  relative to the remaining portion of the inlet lip section  38 , it should be understood that any portion of the inlet lip section  38  may be drooped to improve the aircraft engine performance and reduce nacelle drag at all flight conditions. 
         [0053]    The adaptive inlet lip section  38  improves aerodynamic performance of the gas turbine engine  10  during all operability conditions experienced by the aircraft. In addition, because of the shape changing capabilities of the inlet lip section  38 , the aircraft may be designed having a “thin” inlet lip section  38  (i.e., a slim line nacelle having a reduced contraction ratio is achieved). As a result, the nacelle assembly  26  is designed for specific cruise conditions of the aircraft. A reduced maximum diameter of the nacelle assembly  26  may therefore be achieved while reducing weight, reducing drag, reducing fuel burn and increasing the overall efficiency of the gas turbine engine  10 . 
         [0054]    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 invention.