Abstract:
A propulsion system includes an engine with a fan and a nacelle that circumscribes the engine. The engine defines an axis. A distance between a forwardmost point at the top of the nacelle and at the front of the nacelle to a forwardmost point of the fan is greater than a distance between a forwardmost point at the bottom of the nacelle and at the front of the nacelle.

Description:
CROSS-REFERENCE TO RELATED APPLICATION(S) 
     This application is a divisional of U.S. patent application Ser. No. 12/434,290, filed May 1, 2009, entitled “CAMBERED AERO-ENGINE INLET”. 
    
    
     BACKGROUND 
     The present invention relates generally to gas turbine engines, and more particularly, to turbofan aircraft engines. 
     One of the primary design criteria for aircraft turbofan engines is to propel an aircraft in flight with maximum efficiency, thereby reducing fuel consumption. Thus, turbofan engines are continually being developed and improved to maximize thrust capability with the greatest aerodynamic efficiency possible. 
     One of the components of the turbofan engine is an array of fan blades, which are positioned adjacent the forward portion of the turbofan adjacent the turbofan&#39;s inlet. The fan blades produce thrust, and thus, are typically designed to maximize the aerodynamic loading and the amount of propulsion thrust generated thereby during operation. However, fan loading is limited by stall, flutter, or other instability parameters of the air being pressurized. 
     Fan stall margin is a fundamental design requirement for the turbofan and is affected by aerodynamic fan loading. A major factor affecting the aerodynamic loading of the fan is the geometry of the inlet upstream of the fan. Aircraft wings are known to induce an upwash velocity in airflow in the inlet. Unfortunately, conventional inlets do not account sufficiently for the vertical component (vector) of the airflow. The circumferential vector of the airflow causes the fan to operate with different circumferential sectors having different flow-pressure ratio characteristics. Specifically, the circumferential vector of the inlet airflow manifests itself as a swirl velocity along the face of the fan blades. The swirl velocity is in a direction counter to the direction of rotation of the fan along at least a portion of the fan face. The swirl velocity has a destabilizing effect on the flow condition over sector(s) of the fan, and thus, degrades the stall margin of the fan. 
     Similarly, the different flow-pressure ratio characteristics on the fan produce a circumferential variation in total pressure at the inlet to the core stream of the turbofan. The total pressure variation (distortion) has a destabilizing effect on the operation of the low pressure compressor and high pressure compressor sections of the turbofan. 
     SUMMARY 
     An inlet of a nacelle for channeling inlet airflow to a fan of a gas turbine engine includes a forward lip and an inner surface. The forward lip extends around a centerline axis of the engine and is adapted to accommodate a flow angle of incoming airflow. The inner surface extends from the forward lip along the centerline axis of the engine to adjacent the fan. The inner surface has a first profile above the centerline axis of the engine and a second profile below the centerline axis of the engine. The first profile and the second profile are adapted to define an inlet centerline axis that extends below the centerline axis of the engine at a face of the fan. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG. 1  is a schematic representation of a portion of an aircraft having a wing mounted gas turbine engine including a nacelle inlet. 
         FIG. 2  is a partially schematic partially sectional view of the turbofan portion of the gas turbine engine of  FIG. 1 . 
         FIG. 3  is a schematic sectional view of the nacelle inlet and a fan shown in  FIG. 2 . 
         FIG. 4  is a diagrammatic view of the front of the fan showing a decrease in an airflow swirl velocity. 
     
    
    
     DETAILED DESCRIPTION 
       FIG. 1  shows an exemplary portion of an aircraft  10  having at least one gas turbine engine  12  mounted to one wing  14  thereof. The engine  12  includes a nacelle  16  that defines an inlet  18 . 
     The engine  12  is disposed around an engine centerline axis  20 . The nacelle  16  is an aerodynamic structure that surrounds at least a portion of the engine  12  and forms an inlet  18  thereto. The nacelle  16  channels free stream airflow  22  through the inlet  18  to the internal components of the engine  12 . Adjacent to the inlet  18  and wing  14 , the free stream airflow  22  has an upwash angle β with respect to the engine centerline axis  20 . The upwash angle β is due to the aerodynamic affects of the wing  14 . The upwash angle β exists in the inlet airflow after the free stream airflow  22  enters the inlet  18 . 
     As will be discussed subsequently, the portion of the nacelle  16  that defines the inlet  18  is adapted to align the upwash angle β with the engine centerline axis  20  at a fan within the engine  12 . The alignment of the upwash angle β with the engine centerline axis  20  reduces or eliminates the vertical vector of the inlet airflow that causes a swirl velocity along the face of the fan blades. The reduction or elimination of the swirl velocity allows components of the engine  12  to better meet fan stall margin requirements and improves aerodynamic fan loading. 
       FIG. 2  shows a section of the turbofan portion of the gas turbine engine  12 . The engine  12  includes a fan case  24  having a fan case flange  26 , a fan  28 , a low pressure compressor  30 , a high pressure compressor  32 , a combustor  34 , a high pressure turbine  36 , and a low pressure turbine  38 . The nacelle  16  includes forward lip portion  40 , a crown portion  42 , a keel portion  44  and an inner surface  46 . 
     In  FIG. 2 , a section of the nacelle  16  has been removed along a vertical plane extending through top dead center and bottom dead center thereof. The nacelle  16  connects to and extends around the axisymmetrical fan case  24 . The fan case flange  26  extends generally radially outward from the forward portion of the fan case  24  to connect to the nacelle  16 . The portion of the nacelle  16  that defines the inlet  18  extends forward of the fan case  24  and fan case flange  26 . The fan  28  is rotationally mounted within the fan case  24  and is co-aligned along the same axis as the engine centerline axis  20 . The low pressure compressor  30 , high pressure compressor  32 , combustor  34 , high pressure turbine  36 , and low pressure turbine  38  extend in series axially downstream (rearward) of the fan  28 . 
     The forward lip  40  of the nacelle  16  defines the forward portion of the inlet  18 . The forward lip  40  extends around the engine centerline axis  20 . The nacelle  16  can generally be divided into the crown portion  42 , which extends generally above the engine centerline axis  20 , and the keel portion  44 , which extends generally below the engine centerline axis  20 . The crown portion  42  is asymmetrical with respect to the keel portion  44  but both together integrally extend around the engine centerline axis  20 . The keel  44  can be dropped (i.e. adapted to have its portion of the forward lip  40  be disposed to the rearward of the forward lip  40  of the crown portion  42 ) such that the forward lip  40  does not extend along a vertical plane. This arrangement allows the inlet  18  to accommodate the upwash angle β of incoming airflow. 
     The inner surface  46  of the nacelle  16  defines a portion of the inlet  18  which also extends through the annular fan case  24  to the fan case flange  26 . As will be discussed subsequently, the inner surface  46  of the nacelle  16  is profiled on both the crown portion  42  and the keel portion  44  to align the upwash of inlet airflow with the engine centerline axis  20  at the face of the fan  28 . 
     During operation of the engine  12 , the free stream airflow  22  enters the inlet  18 , is aligned by the geometry of the nacelle  16 , and is pressurized by the fan  28 . A portion of the airflow is channeled to engine core where it is pressurized, mixed with fuel/combusted, and expanded before being expelled from the engine  12 . A second portion of the airflow bypasses the engine core and is expelled from an outlet portion of the nacelle  16 . 
     After being pressurized in the low pressure compressor  30  and high pressure compressor  32 , the air is mixed with fuel in the combustor  34  for generating hot combustion gases. From the combustor  32 , the gases and airflow are discharged downstream into the high pressure turbine  34 . The high pressure turbine  34  and low pressure turbine  36  in turn receive the combusted gases and extract energy therefrom. The high pressure turbine  34  is joined by a rotor or shaft (not shown) to the high pressure compressor  32  and the fan  28  for powering these components during operation. The combination of thrust produced from the fan  28  and the components of the engine core propel the aircraft  10  in flight ( FIG. 1 ). 
       FIG. 3  shows a schematic sectional view of one embodiment of the nacelle  16  and the fan  28 . The axially forward most extent of the fan case flange  26  is denoted by a fan case flange plane  48 . Similarly, the axial forward most portion of the fan  28  is indicated by a fan face plane  50 . The inner surface  46  of the nacelle  16  includes a first profile  52  and a second profile  54 . The first profile  52  has a generally convex forward section  52 F that transitions to a generally concave rearward section  52 R (when observed from the engine centerline axis  20 ) at inflection point  53 . Similarly, the second profile  54  has a generally convex forward section  54 F that transitions to a generally concave rearward section  52 R (when viewed from the engine centerline axis  20 ) at inflection point  55 . The geometry of the inner surface  46  of the nacelle  16  defines an inlet centerline axis  56 . 
     The forward lip  40  of the nacelle  16  defines the orifice of the inlet  18 . The forward lip  40  extends around the engine centerline axis  20 . The crown portion  42  of the nacelle  16  extends above the engine centerline axis  20 . The keel portion  44  of the nacelle  16  extends below the engine centerline axis  20 . Together both the crown portion  42  and the keel portion  44  form the inner surface  46 . The inner surface  46  defines the inlet  18 , which extends axially along the engine centerline axis  20  from the forward lip  40  to the fan case flange plane  48 . The inner surface  46  also defines the inlet centerline axis  56  which approximates the geometric center of inlet  18 . 
     The shape of the inner surface  46  above the engine centerline axis  20  is asymmetrical with respect to the shape of the inner surface  46  below the engine centerline axis  20 . More particularly, the inner surface  46  above the engine centerline axis  20  is shaped with the first profile  52 , which extends generally axially along the engine centerline axis  20  from the forward lip  40  to at least the fan case flange plane  48 . The first profile  52  varies in the radial distance it is disposed from the engine centerline axis  20 . Similarly, the inner surface  46  below the engine centerline axis  20  is shaped with the second profile  54 , which extends generally axially along the engine centerline axis  20  from the forward lip  40  to at least the fan case flange plane  48 . The second profile  54  varies in the radial distance it is disposed from the engine centerline axis  20 . 
     For most of the axial travel of the first profile  52  and second profile  54  along the engine centerline axis  20 , the radial distance the first profile  52  is disposed from the engine centerline axis  20  differs from the radial distance the second profile  54  is disposed from the engine centerline axis  20 . Thus, the volume and cross sectional area of the inlet  18  above the engine centerline axis  20  generally differs from the volume and cross sectional area of the inlet  18  below the engine centerline axis  20 . The difference in geometry between the first and second profiles  52  and  54  is approximated by the inlet centerline axis  56  (which approximates the geometric center of the inlet  18 ), which has a cambered shape with respect to the engine centerline axis  20 . 
     The inflection point  53  of the first profile  52  is radially and axially offset along the engine centerline axis  20  from the inflection point  55  of the second profile  54 . More particularly, the convex forward section  54 F (when viewed from the engine centerline axis  20 ) of the second profile  54  extends further rearward toward the fan face plane  50  than does the convex forward section  52 F of the first profile  52 . Terms such as “forward,” “rearward,” “upstream,” and “downstream” are defined by the direction of airflow  22  within the inlet  18 . Thus, the concave rearward section  52 R (when viewed from the engine centerline axis  20 ) of the first profile  52  extends further forward away from the fan face plane  50  than does the concave rearward section  54 R of the second profile  54 . This geometry disposes the inflection point  55  between the convex and concave sections ( 54 F and  54 R) of the second profile  54  closer to the face of the fan  28  than the inflection point  53  between the convex and concave sections ( 52 F and  54 R) of the first profile  52 . 
     In particular, the convex forward section  54 F of the second profile  54  is geometrically accentuated relative to the concave rearward section  54 R of the second profile  54  and the convex forward section  52 F of the first profile  52 . Similarly, the concave rearward section  52 R of the first profile  52  is geometrically accentuated relative to the convex forward section  52 F of the first profile  52  and the concave rearward section  54 R or the second profile  54 . The precise interrelation of profiles  52  and  54  (i.e. the precise geometric accentuation of the concave/convex sections  52 F,  52 R,  54 F,  54 R relative to one another) is achieved through computational fluid mechanics. The different axial and radial geometry of the inner surface  46  above the engine centerline axis  20  with respect to the axial and radial geometry of the inner surface  46  below the engine centerline axis  20  cambers the inlet centerline axis  56  with respect to the engine centerline axis  20 . 
     In one embodiment, the geometry of the first profile  52  with respect to the second profile  54  disposes the inlet centerline axis  56  radially above the engine centerline axis  20  immediately to the rear of the forward lip  40  within the inlet  18 . The inlet centerline axis  56  extends above the engine centerline axis  20  for a portion of its axial extent within the inlet  18 . In one embodiment, the geometry of the first profile  52  with respect to the second profile  54  about the engine centerline axis  20  causes the inlet centerline axis  56  to have an acute descending depression angle a at the fan case flange plane  48 . In one embodiment, the acute depression angle a the inlet centerline axis  56  forms with respect to the engine centerline axis  20  offsets any residual upwash of the airflow  22  entering the inlet  18 . 
     The geometry of the inner surface  46  allows the airflow  22  in the inlet  18  to substantially align with the engine centerline axis  20  (which is also the rotational axis of the fan  28 ) at the face of the fan  28 . In the embodiment shown, the inlet centerline axis  56  coincident with the engine centerline axis  20  at the fan case flange plane  48 . In other embodiments, the inlet centerline axis  56  can merge with or be coincident with the engine centerline axis  20  forward (as defined by the direction of airflow  22 ) of the fan case flange plane  48 . 
     As illustrated in  FIG. 4 , an airflow  22  swirl velocity on the face of the fan  28  at 90° and 270° to top dead center (TDC) is reduced or eliminated by offsetting the first profile  52  with respect to the second profile  54 . The elimination of the swirl velocity occurs because the inlet centerline axis  56  extends below the engine centerline axis  20  and aligns the airflow  22  in the inlet  18  with the engine centerline axis  20  at the face of the fan  28 . By aligning the airflow  22  in the inlet  18  generally with the engine centerline axis  20 , the vertical vector of the airflow  22  that causes the airflow  22  swirl velocity is reduced or eliminated. Thus, the swirl velocity at 270° (measured from TDC of the fan  28 ), co-rotating with the fan  28  (when the fan  28  is rotating in a clockwise direction), and the swirl velocity at 90° counter-rotating with the fan  28  (when the fan  28  is rotating in a clockwise direction) are reduced or eliminated. The reduction or elimination of the airflow  22  swirl velocity allows the components of the engine  12  to better meet fan stall margin requirements and improves aerodynamic loading. 
     Although the present invention has been described with reference to preferred embodiments, workers skilled in the art will recognize that changes may be made in form and detail without departing from the spirit and scope of the invention.