Abstract:
A gas turbine engine system according to an exemplary aspect of the present disclosure may include a core engine defined about an axis, a fan driven by the core engine about the axis to generate bypass flow, and at least one integrated mechanism in communication with the bypass flow. The bypass flow defines a bypass ratio greater than about six (6). The at least one integrated mechanism includes a variable area fan nozzle (VAFN) and thrust reverser, and a plurality of positions to control bypass flow.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
       [0001]    This application is a continuation-in-part of U.S. application Ser. No. 12/440,746 filed Mar. 11, 2009, which is a National Phase application of PCT/US2006/039990 filed Oct. 12, 2006. 
     
    
     BACKGROUND 
       [0002]    This invention relates to gas turbine engines and, more particularly, to a gas turbine engine having a variable fan nozzle integrated with a thrust reverser of the gas turbine engine. 
         [0003]    Gas turbine engines are widely known and used for power generation and vehicle (e.g., aircraft) propulsion. A typical gas turbine engine includes a compression section, a combustion section, and a turbine section that utilize a primary airflow into the engine to generate power or propel the vehicle. The gas turbine engine is typically mounted within a housing, such as a nacelle. A bypass airflow flows through a passage between the housing and the engine and exits from the engine at an outlet. 
         [0004]    Presently, conventional thrust reversers are used to generate a reverse thrust force to slow forward movement of a vehicle, such as an aircraft. One type of conventional thrust reverser utilizes a moveable door stowed near the rear of the nacelle. After touch-down of the aircraft for landing, the door moves into the bypass airflow passage to deflect the bypass airflow radially outwards into cascades, or vents, that direct the discharge airflow in a forward direction to slow the aircraft. Although effective, this and other conventional thrust reversers serve only for thrust reversal and, when in the stowed position for non-landing conditions, do not provide additional functionality. The use of a variable area fan nozzle (VAFN) has been proposed for low pressure ratio fan designs to improve the propulsive efficiency of high bypass ratio gas turbine engines. Integrating the VAFN functionality into a common set of thrust reverser cascades operated by a common actuation system represents a significant reduction in complexity and weight. 
       SUMMARY 
       [0005]    A gas turbine engine system according to an exemplary aspect of the present disclosure may include a core engine defined about an axis, a fan driven by the core engine about the axis to generate bypass flow, and at least one integrated mechanism in communication with the bypass flow. The bypass flow defines a bypass ratio greater than about six (6). The at least one integrated mechanism includes a variable area fan nozzle (VAFN) and thrust reverser, and a plurality of positions to control bypass flow. 
         [0006]    In a further non-limiting embodiment of any of the foregoing gas turbine engine system embodiments, the bypass flow is arranged to communicate with an exterior environment when the integrated mechanism is in a deployed position. 
         [0007]    In a further non-limiting embodiment of any of the foregoing gas turbine engine system embodiments, the integrated mechanism includes a plurality of apertures to enable the communication of the bypass flow with the exterior environment when the integrated mechanism is in the deployed position. 
         [0008]    In a further non-limiting embodiment of any of the foregoing gas turbine engine system embodiments, the integrated mechanism includes a single actuator set to move between the plurality of positions. 
         [0009]    In a further non-limiting embodiment of any of the foregoing gas turbine engine system embodiments, the thrust reverser has a stowed position and a deployed position to divert the bypass flow in a thrust reversing direction. 
         [0010]    In a further non-limiting embodiment of any of the foregoing gas turbine engine system embodiments, a gear system is driven by the core engine. The fan is driven by the gear system. The gear system defines a gear reduction ratio of greater than about 2.3. 
         [0011]    In a further non-limiting embodiment of any of the foregoing gas turbine engine system embodiments, a gear system is driven by the core engine. The fan is driven by the gear system. The gear system defines a gear reduction ratio of greater than 2.5. 
         [0012]    In a further non-limiting embodiment of any of the foregoing gas turbine engine system embodiments, the core engine includes a low pressure turbine which defines a pressure ratio that is greater than about five (5). 
         [0013]    In a further non-limiting embodiment of any of the foregoing gas turbine engine system embodiments, the core engine includes a low pressure turbine which defines a pressure ratio that is greater than five (5). 
         [0014]    In a further non-limiting embodiment of any of the foregoing gas turbine engine system embodiments, the at least one integrated mechanism is arranged to change a pressure ratio across the fan. 
         [0015]    In a further non-limiting embodiment of any of the foregoing gas turbine engine system embodiments, the bypass ratio is greater than about 10. 
         [0016]    In a further non-limiting embodiment of any of the foregoing gas turbine engine system embodiments, the bypass ratio is greater than 10. 
         [0017]    In a further non-limiting embodiment of any of the foregoing gas turbine engine system embodiments, a gear system is driven by the core engine. The fan is driven by the gear system with a gear reduction ratio greater than 2.5. The gear system is an epicycle gear train. The core engine includes a low pressure turbine which defines a pressure ratio that is greater than five (5). 
         [0018]    A gas turbine engine according to another exemplary aspect of the present disclosure may include a core engine defined about an axis, a fan couple to be driven by said core engine about the axis to generate a bypass flow, and at least one integrated mechanism in communication with the bypass flow. The core engine includes at least a low pressure turbine which defines a pressure ratio that is greater than about five (5). The at least one integrated mechanism includes a variable area fan nozzle (VAFN) and a thrust reverser. The integrated mechanism also may includes a plurality of positions to control bypass flow. The integrated mechanism includes a section common to the thrust reverser and VAFN. 
         [0019]    In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the integrated mechanism includes at least one actuator set to move between the plurality of positions. 
         [0020]    In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the thrust reverser has a stowed position and a deployed position to divert the bypass flow in a thrust reversing direction. 
         [0021]    In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the common section is moveable between a plurality of axial positions and has a plurality of apertures providing a flow path for the bypass flow to reach an exterior environment of the gas turbine engine. 
         [0022]    In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, a gear system is included. The core engine drives the fan via the gear system, which defines a gear reduction ratio of greater than about 2.3. 
         [0023]    In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, a gear system is included. The core engine drives the fan via the gear system, which defines a gear reduction ratio of greater than 2.5. 
         [0024]    In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the bypass flow defines a bypass ratio greater than about ten (10). 
         [0025]    In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the bypass flow defines a bypass ratio greater than ten (10). 
         [0026]    In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the thrust reverser includes a blocker door moveable between a stowed position and a deployed position and a link having one end connected to the blocker door and an opposite end connected to a support. 
         [0027]    In a further non-limiting embodiment of any of the foregoing gas turbine engine embodiments, the blocker door includes a slot having a T-shaped cross section, the slot slidably receiving the link. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0028]    The various features and advantages of this invention will become apparent to those skilled in the art from the following detailed description of the currently preferred embodiment. The drawings that accompany the detailed description can be briefly described as follows. 
           [0029]      FIG. 1  illustrates selected portions of an example gas turbine engine system having a mechanism that integrates a variable fan nozzle integrated and a thrust reverser. 
           [0030]      FIG. 2  illustrates a perspective view of the example gas turbine engine system with cascades exposed for thrust reversal. 
           [0031]      FIG. 3A  illustrates a schematic view of the mechanism having an axially moveable section that is in a closed position. 
           [0032]      FIG. 3B  illustrates a schematic view of the axially moveable section in an intermediate position for altering a discharge flow from the gas turbine engine. 
           [0033]      FIG. 3C  illustrates a schematic view of the axially moveable section in an open position for generating a thrust reversing force. 
           [0034]      FIG. 4  illustrates a blocker door of the thrust reverser. 
           [0035]      FIG. 5  illustrates a view of an example slot of the blocker door according to the section shown in  FIG. 4 . 
       
    
    
     DETAILED DESCRIPTION 
       [0036]      FIG. 1  illustrates a schematic view of selected portions of an example gas turbine engine  10  suspended from an engine pylon  12  of an aircraft, as is typical of an aircraft designed for subsonic operation. The gas turbine engine  10  is circumferentially disposed about an engine centerline, or axial centerline axis A. The gas turbine engine  10  includes a fan  14 , a low pressure compressor  16   a , a high pressure compressor  16   b , a combustion section  18 , a low pressure turbine  20   a , and a high pressure turbine  20   b . As is well known in the art, air compressed in the compressors  16   a ,  16   b  is mixed with fuel that is burned in the combustion section  18  and expanded in the turbines  20   a  and  20   b . The turbines  20   a  and  20   b  are coupled for rotation with, respectively, rotors  22  and  24  (e.g., spools) to rotationally drive the compressors  16   a ,  16   b  and the fan  14  in response to the expansion. In this example, the rotor  22  also drives the fan  14  through a gear train  24 . 
         [0037]    The engine  10  is preferably a high-bypass geared architecture aircraft engine. In one disclosed, non-limiting embodiment, the engine  10  bypass ratio is greater than about six (6) to ten (10), the gear train  22  is an epicyclic gear train such as a planetary gear system or other gear system with a gear reduction ratio of greater than about 2.3 and the low pressure turbine  18  has a pressure ratio that is greater than about 5. In the example shown, the gas turbine engine  10  is a high bypass turbofan arrangement. In one example, the bypass ratio is greater than 10, and the fan  14  diameter is substantially larger than the diameter of the low pressure compressor  16   a . The low pressure turbine  20   a  has a pressure ratio that is greater than 5, in one example. The gear train  24  is an epicycle gear train, for example, a star gear train, providing a gear reduction ratio of greater than 2.5. 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 engines. 
         [0038]    An outer housing, nacelle  28 , (also commonly referred to as a fan nacelle) extends circumferentially about the fan  14 . A fan bypass passage  32  extends between the nacelle  28  and an inner housing, inner cowl  34 , which generally surrounds the compressors  16   a ,  16   b  and turbines  20   a ,  20   b . In this example, the gas turbine engine  10  includes integrated mechanisms  30  that are coupled to the nacelle  28 . The integrated mechanisms  30  integrate functions of a variable fan nozzle and a thrust reverser, as will be described below. Any number of integrated mechanisms  30  may be used to meet the particular needs of an engine. In this example, two integrated mechanisms  30  are used, one on each semi-circular half of the nacelle  28 . 
         [0039]    In operation, the fan  14  draws air into the gas turbine engine  10  as a core flow, C, and into the bypass passage  32  as a bypass air flow, D. The bypass air flow D is discharged as a discharge flow through a rear exhaust  36  associated with the integrated mechanism  30  near the rear of the nacelle  28  in this example. The core flow C is discharged from a passage between the inner cowl  34  and a tail cone  38 . 
         [0040]    For the gas turbine engine  10  shown  FIG. 1 , a significant amount of thrust may be provided by the discharge flow due to the high bypass ratio. 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 or to enhance conditions for aircraft control, operation of the fan  14 , operation of other components associated with the bypass passage  32 , or operation of the gas turbine engine  10 . For example, an effective reduction in area of the rear exhaust  36  causes an air pressure increase within the bypass passage  32  that in turn changes a pressure ratio across the fan  14 . 
         [0041]    In the disclosed example, the integrated mechanism  30  includes a structure associated with the rear exhaust  36  to change one or more of these parameters. However, it should be understood that the bypass flow or discharge flow may be effectively altered by other than structural changes, for example, by altering a flow boundary layer. Furthermore, it should be understood that effectively altering a cross-sectional area of the rear exhaust  36  is not limited to physical locations approximate to the exit of the nacelle  28 , but rather, includes altering the bypass flow D by any suitable means. 
         [0042]    Referring to  FIGS. 2 and 3A , the integrated mechanism  30  in this example includes a nozzle  40  and a thrust reverser  42 . The nozzle  40  and thrust reverser include a common part, section  44 , which is moveable between a plurality of axial positions relative to the centerline axis A. In this example, the section  44  is a hollow sleeve-like structure that extends about a cascade section  46 . Actuators  48  are mounted within the nacelle  28  in this example. Links  50  extend through the cascade section  46  and are coupled on one end with the respective actuators  48  and on an opposite end with the section  44  in a known manner. A controller  49  communicates with the actuators  48  to selectively axially move the section  44 . The controller  49  may be dedicated to controlling the integrated mechanism  30 , integrated into an existing engine controller within the gas turbine engine  10 , or be incorporated with other known aircraft or engine controls. Alternatively, one or more of the actuators  48  are mounted within the cascade section  46  in a known manner. 
         [0043]    In the disclosed example, the cascade section  46  includes a plurality of apertures  52 , or vents, that provide a flow path between the bypass passage  32  and the exterior environment of the gas turbine engine  10 . The apertures  52  may be formed in any known suitable shape, such as with airfoil shaped vanes between the apertures. In this example, the apertures  52  are arranged in circumferential rows about the cascade section  46 . A first set of apertures  52   a  near the forward end of the cascade section  46  are angled aft and a second set of apertures  52   b  aft of the first set of apertures  52   a  are angled forward. Axial movement of the section  44  selectively opens, or exposes, the apertures  52   a , apertures  52   b , or both to provide an auxiliary passage for the discharge flow, as will be described below. 
         [0044]    In the illustrated example, there are two circumferential rows in the first set of apertures  52   a  and a larger number of circumferential rows in the second set of apertures  52   b . In one example, two circumferential rows in the first set of apertures  52   a  is adequate for altering the discharge flow, as will be described. However, it is to be understood that one circumferential row or greater than two circumferential rows may be used for smaller or larger alterations, respectively. 
         [0045]    The thrust reverser  42  includes a blocker door  62  having a stowed position ( FIG. 3A ) and a fully deployed position ( FIG. 3C ). The blocker door  62  is pivotally connected to the section  44  at connection  63 . A drag link  64  includes one end that is slidably connected to the blocker door  62  and an opposite end that is connected to a support, the inner cowl  34  in this example. Although only one drag link  64  is shown, it is to be understood that any suitable number of drag links  64  may be used. 
         [0046]    Referring to  FIGS. 4 and 5 , the blocker door  62  includes a slot  66  for slidably connecting the drag link  64  to the blocker door  62 . In this example, the shape of the slot  66  is adapted to receive and retain the end of the drag link  64 . For example, the slot  66  is T-shaped and the end of drag link  64  includes laterally extending slide members  68 , such as rollers, bearings, friction material, or other known suitable mechanism for allowing the end of the drag link  64  to slide along the slot  66 . Given this description, one of ordinary skill in the art will recognize alternative suitable slot shapes or sliding connections to meet their particular needs. 
         [0047]    In operation, the controller  49  selectively commands the actuators  48  to move the section  44  between the plurality of axial positions to alter the discharge flow or provide thrust reversal.  FIG. 3A  illustrates the section  44  in a first axial position (i.e., a closed position) sealed against the nacelle  28 . In the closed position, the section  44  completely covers the cascade section  46  such that the discharge flow exits axially through the rear exhaust  36 . 
         [0048]      FIG. 3B  illustrates the section  44  in a second axial position spaced apart from the nacelle  28  to provide an opening there between and expose a portion of the cascade section  46 . In the second position, the first set of apertures  52   a  are exposed to provide an auxiliary passage for the discharge flow. The auxiliary passage provides an additional passage (i.e., additional effective cross-sectional flow area) for exit of the discharge flow from the bypass passage  32  to thereby alter the discharge flow. A portion of the discharge flow flows through the first set of apertures  52   a  and is directed in the aft direction. Although the aft angle in the illustrated example is not parallel to the centerline axis A, a geometric component of the aft angle is parallel. The geometric component of the discharge flow that is parallel to the centerline axis A provides the benefit of maintaining a portion of the thrust generated by the discharge flow. 
         [0049]    Upon movement of the section  44  between the first position and the second position, the blocker door  62  remains in the stowed position. The connection between the drag link  64  and the slot  66  provides a range of lost motion movement. That is, the movement of the section  44  causes the drag link  64  to slide along the slot  66  of the blocker door  62  without moving the blocker door  62  into the deployed position. 
         [0050]      FIG. 3C  illustrates the section  44  in a third axial position (i.e., a thrust reverse position). Movement of the section  44  beyond the second position toward the third position causes the end of the drag link  64  to engage an end  70  of the slot  66 . Once engaged, the drag link  64  pivots the blocker door  62  about the connection  63  and into the bypass passage  32 . The blocker door  62  deflects the discharge flow radially outwards relative to the centerline axis A toward the cascade section  46 . The movement of the section  44  to the third position also exposes the apertures  52   b . The deflected discharge flow enters the second set of apertures  52   b , which angle the discharge flow in the forward direction to generate a reverse thrust force. 
         [0051]    In this example, there are more apertures  52  within the first set of apertures  52   b  than in the second set of apertures  52   a . Thus, the reverse thrust force due to discharge flow through the second set of apertures  52   b  overcomes any thrust due to aft discharge flow from the apertures  52   a.    
         [0052]    The disclosed example integrated mechanism  30  thereby integrates the function of altering the discharge flow with the thrust reversing function. The integrated mechanism  30  utilizes a single set or system of actuators  48  to eliminate the need for separate actuators or sets of actuators for altering the discharge flow and deploying the thrust reverser. Using a single actuator or set of actuators  48  as in the disclosed examples eliminates at least some of the actuators that would otherwise be used, thereby reducing the weight of the gas turbine engine  10  and increasing the fuel efficiency. 
         [0053]    Although a preferred embodiment of this invention has been disclosed, a worker of ordinary skill in this 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.