Patent Publication Number: US-11639686-B2

Title: Aircraft propulsion system with variable area inlet assembly

Description:
BACKGROUND 
     1. Technical Field 
     This disclosure relates generally to an aircraft propulsion system and, more particularly, to an inlet assembly for the aircraft propulsion system with a variable airflow inlet area. 
     2. Background Information 
     An aircraft propulsion system includes an inlet structure and a gas turbine engine. The inlet structure directs air into the gas turbine engine. Some known inlet structures include a variable airflow inlet area for tailoring a mass flow of the air entering the gas turbine engine. While these known inlet structures have various advantages, there is still room in the art for improvement. There is a need in the art therefore for an improved inlet assembly with a variable airflow inlet area. 
     SUMMARY OF THE DISCLOSURE 
     According to an aspect of the present disclosure, an assembly is provided for an aircraft propulsion system. This assembly includes a center body, a scarfed inlet structure and an inlet passage. The scarfed inlet structure extends circumferentially about the center body. The inlet passage is radially between and formed by at least the center body and the scarfed inlet structure. The inlet passage includes a metering portion. A first component of the assembly is configured to rotate about an axis relative to a second component of the assembly between: (A) a first position where the metering portion has a first area; and (B) a second position where the metering portion has a second area that is different than the first area. The first component is configured as or otherwise includes one of the center body or the scarfed inlet structure. The second component is configured as or otherwise includes the other one of the center body or the scarfed inlet structure. 
     According to another aspect of the present disclosure, another assembly is provided for an aircraft propulsion system. This assembly includes a center body, a nacelle inlet structure and an inlet passage. The center body includes a center body surface with a trailing edge. The nacelle inlet structure extends circumferentially around the center body. The nacelle inlet structure has a leading edge. The inlet passage is radially between and formed by at least the center body and the nacelle inlet structure. The inlet passage includes a metering portion. A first component of the assembly is configured to rotate about an axis relative to a second component of the assembly between: (A) a first position where the metering portion has a first area, and where the trailing edge is axially displaced from the leading edge by a first distance; and (B) a second position where the metering portion has a second area that is different than the first area, and where the trailing edge is axially displaced from the leading edge by a second distance that is different than the first distance. The first component of the assembly is configured as or otherwise includes a first of the center body or the nacelle inlet structure. The second component of the assembly is configured as or otherwise includes a second of the center body or the nacelle inlet structure. 
     According to still another aspect of the present disclosure, another assembly is provided for an aircraft propulsion system. This assembly includes a double tapered center body, a nacelle inlet structure and an inlet passage. The nacelle inlet structure extends circumferentially around the double tapered center body. The inlet passage is radially between and formed by at least the double tapered center body and the nacelle inlet structure. The inlet passage includes a metering portion. A first component of the assembly is configured to rotate about an axis relative to a second component of the assembly between: (A) a first position where the metering portion has a first area; and (B) a second position where the metering portion has a second area that is different than the first area. The first component of the assembly is configured as or otherwise includes one of the double tapered center body or the nacelle inlet structure. The second component of the assembly is configured as or otherwise includes the other one of the double tapered center body or the nacelle inlet structure. 
     The nacelle inlet structure may be configured as or otherwise include as a scarfed inlet structure. 
     The nacelle inlet structure may be configured as or otherwise include a scarfed inlet lip. 
     The center body may also include a first tubular surface adjacent the center body surface. The first tubular surface may taper radially inwards towards the axis as the center body extends in a first direction along the axis away from the center body surface. 
     The center body may also include a second tubular surface adjacent the center body surface. The second tubular surface may taper radially inwards towards the axis as the center body extends in a second direction along the axis away from the center body surface. 
     The first component of the assembly may be configured as or otherwise include the center body. The second component of the assembly may be configured as or otherwise include the scarfed inlet structure. 
     The scarfed inlet structure may have a leading edge. A first point on the leading edge may be axially displaced from a second point on the leading edge along the axis. 
     A plane of an inlet orifice to the inlet opening passage may be angularly offset from the axis. 
     The center body may include a tubular surface that radially tapers as center body extends in a forward direction along the axis. 
     The center body may include a tubular surface that radially tapers as center body extends in an aft direction along the axis. 
     The center body may include a first tapered surface, a second tapered surface and a plateau surface. The plateau surface may extend axially between the first tapered surface and the second tapered surface. 
     The metering portion may have an inner peripheral boundary. The inner peripheral boundary may be formed by the plateau surface when the first component of the assembly is in the first position. The inner peripheral boundary may be formed by the plateau surface and the second tapered surface when the first component of the assembly is in the second position. 
     The center body may include a plateau surface aligned with an inlet lip of the scarfed inlet structure. A trailing edge of the plateau surface may have a scarfed configuration. 
     The center body may include a plateau surface with a trailing edge. A first point on the trailing edge may be axially displaced from a second point on the trailing edge along the axis. 
     The scarfed inlet structure may include a leading edge. The center body may include a plateau surface with a trailing edge. The trailing edge may be separated from the leading edge by a first axial distance along the axis when the first component of the assembly is in the first position. The trailing edge may be separated from the leading edge by a second axial distance along the axis when the first component of the assembly is in the second position. The second axial distance may be different than the first axial distance. 
     The center body may include a cylindrical surface with an axial length that changes as the cylindrical surface extends circumferentially about the axis. 
     The axis may be coaxial with a centerline of the aircraft propulsion system. 
     The axis may be offset from a centerline of the aircraft propulsion system. 
     The present disclosure may include any one or more of the individual features disclosed above and/or below alone or in any combination thereof. 
     The foregoing features and the operation of the invention will become more apparent in light of the following description and the accompanying drawings. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG.  1    is a schematic side sectional illustration of an aircraft propulsion system with a variable area inlet assembly in a first position. 
         FIG.  2    is a cross-sectional illustration of the variable area inlet assembly in the first position taken along line  2 - 2  in  FIG.  1   . 
         FIG.  3    is a cross-sectional illustration of the variable area inlet assembly in the first position taken along line  3 - 3  in  FIG.  1   . 
         FIG.  4    is a cross-sectional illustration of the variable area inlet assembly in a second position taken along line  4 - 4  in  FIG.  1   . 
         FIG.  5    is a side sectional illustration of a forward, upstream portion of a nacelle inlet structure of the variable area inlet assembly. 
         FIG.  6    is a side sectional illustration of a center body of the nacelle inlet structure. 
         FIG.  7    is a side sectional illustration of the variable area inlet assembly in the first position. 
         FIG.  8    is a side sectional illustration of the variable area inlet assembly in the second position. 
         FIG.  9    is a side sectional illustration of another variable area inlet assembly. 
     
    
    
     DETAILED DESCRIPTION 
       FIG.  1    illustrates an aircraft propulsion system  20  for an aircraft such as, but not limited to, a commercial airliner or cargo plane. The aircraft propulsion system  20  includes a gas turbine engine  22  and a nacelle  24 . 
     The gas turbine engine  22  may be configured as a high-bypass turbofan engine. The gas turbine engine  22  of  FIG.  1   , for example, includes a fan section  26 , a compressor section  27 , a combustor section  28  and a turbine section  29 . The compressor section  27  may include a low pressure compressor (LPC) section  27 A and a high pressure compressor (HPC) section  27 B. The turbine section  29  may include a high pressure turbine (HPT) section  29 A and a low pressure turbine (LPT) section  29 B. 
     The engine sections  26 - 29 B are arranged sequentially along an axial centerline  30  (e.g., a rotational axis) of the gas turbine engine  22  within an aircraft propulsion system housing  32 . This housing  32  includes an outer housing structure  34  and an inner housing structure  36 . 
     The outer housing structure  34  includes an outer case  38  (e.g., a fan case) and an outer structure  40  of the nacelle  24 ; i.e., an outer nacelle structure. The outer case  38  houses at least the fan section  26 . The outer nacelle structure  40  houses and provides an aerodynamic cover for the outer case  38 . The outer nacelle structure  40  also covers a portion of an inner structure  42  of the nacelle  24 ; i.e., an inner nacelle structure, which may also be referred to as an inner fixed structure (IFS). More particularly, the outer nacelle structure  40  axially overlaps and extends circumferentially about (e.g., completely around) the inner nacelle structure  42 . The outer nacelle structure  40  and the inner nacelle structure  42  thereby at least partially or completely form a bypass flow path  44  within the aircraft propulsion system  20 . 
     The inner housing structure  36  includes an inner case  46  (e.g., a core case) and the inner nacelle structure  42 . The inner case  46  houses one or more of the engine sections  27 A- 29 B, which engine sections  27 A- 29 B may be collectively referred to as an engine core. The inner nacelle structure  42  houses and provides an aerodynamic cover for the inner case  46 . 
     Each of the engine sections  26 ,  27 A,  27 B,  29 A and  29 B includes a bladed rotor  48 - 52 . The fan rotor  48  and the LPC rotor  49  are connected to and driven by the LPT rotor  52  through a low speed shaft  54 . The HPC rotor  50  is connected to and driven by the HPT rotor  51  through a high speed shaft  56 . The shafts  54  and  56  are rotatably supported by a plurality of bearings (not shown). Each of these bearings is connected to the aircraft propulsion system housing  32  (e.g., the inner case  46 ) by at least one stationary structure such as, for example, an annular support strut. 
     During operation, air enters the aircraft propulsion system  20  through an inlet structure  58  of the outer nacelle structure  40 ; i.e., a nacelle inlet structure. This air is directed through a duct  60  (e.g., a fan duct in the fan section  26 ) and into a core flow path  62  and the bypass flow path  44 . The core flow path  62  extends axially along the axial centerline  30  within the aircraft propulsion system  20 , through the engine sections  27 A- 29 B, to a core nozzle outlet, where the core flow path  62  is radially within the inner case  46 . The bypass flow path  44  extends axially along the axial centerline  30  within the aircraft propulsion system  20  to a bypass nozzle outlet, where the bypass flow path  44  is radially between the outer nacelle structure  40  and the inner nacelle structure  42 . The air within the core flow path  62  may be referred to as “core air”. The air within the bypass flow path  44  may be referred to as “bypass air”. 
     The core air is compressed by the compressor rotors  49  and  50  and directed into a combustion chamber of a combustor in the combustor section  28 . Fuel is injected into the combustion chamber and mixed with the compressed core air to provide a fuel-air mixture. This fuel air mixture is ignited and combustion products thereof flow through and sequentially cause the turbine rotors  51  and  52  to rotate. The rotation of the turbine rotors  51  and  52  respectively drive rotation of the compressor rotors  50  and  49  and, thus, compression of the air received from a core airflow inlet. The rotation of the LPT rotor  52  also drives rotation of the fan rotor  48 , which propels bypass air through and out of the bypass flow path  44 . The propulsion of the bypass air may account for a majority of thrust generated by the turbine engine  22 . The aircraft propulsion system  20  of the present disclosure, however, is not limited to the exemplary gas turbine engine configuration described above. 
     Optimal mass flow requirements of the air entering the aircraft propulsion system  20  through the nacelle inlet structure  58  may change depending upon one or more parameters. These parameters may include, but are not limited to, modes of operation, aircraft maneuvers and operating conditions. For example, where the aircraft flies at supersonic speeds, the nacelle inlet structure  58  may be configured to direct a first mass flow of the air into the aircraft propulsion system  20 . When the aircraft flies at subsonic speeds, the nacelle inlet structure  58  may be configured to direct a second mass flow of the air into the aircraft propulsion system  20 , where the second mass flow is greater than the first mass flow. 
     To accommodate changing mass flows, the aircraft propulsion system  20  of  FIG.  1    includes a variable area inlet assembly  64 . The variable area inlet assembly  64  of  FIGS.  1  and  2    includes at least the nacelle inlet structure  58  and a center body  66 ; e.g., an inlet cone or an inlet spike. At least these variable area inlet assembly components  58  and  66  collectively form an annular inlet duct, which inlet duct forms an inlet passage  68  into and within the aircraft propulsion system  20 . Referring to  FIG.  1   , the inlet duct and its inlet passage  68  extend axially along the axial centerline  30  from an inlet orifice  70  of the inlet passage  68  to the fan duct  60 . The inlet passage  68  of  FIG.  1    includes a metering portion  72 ; e.g., a choke point. The term “metering portion” may describe a portion of the inlet passage  68  with the smallest cross-sectional flow area. 
     The variable area inlet assembly  64  is configured to provide a variable airflow inlet area. More particularly, at least the variable area inlet assembly components  58  and  66  are configured to provide the metering portion  72  of the inlet passage  68  with a variable cross-sectional area. Referring to  FIG.  3    for example, during a first (e.g., supersonic) mode of operation, at least one of the variable area inlet assembly components  58  and  66  is configured in/moves (e.g., rotates) to a first position; e.g., a fully closed position. In this first position, the metering portion  72  has a first area  74 A; e.g., a cross-sectional area in the plane of  FIG.  3   . Referring to  FIG.  4   , during a second (e.g., subsonic) mode of operation, the at least one of the variable area inlet assembly components  58  and  66  is configured in/moves (e.g., rotates) to a second position; e.g., a fully open position. In this second position, the metering portion  72  has a second area  74 B; e.g., a cross-sectional area in the plane of  FIG.  4   . The second area  74 B is different (e.g., greater) than the first area  74 A. 
     Referring to  FIG.  1   , the nacelle inlet structure  58  is disposed at a forward, upstream end of the nacelle  24 . The nacelle inlet structure  58  may be configured as a stationary and/or scarfed inlet structure. The nacelle inlet structure  58  of  FIG.  1    includes a tubular inner barrel  76 , a tubular outer barrel  78  and an annular scarfed inlet lip  80 . 
     The inner barrel  76  extends circumferentially about (e.g., completely around) an axial centerline  82  (e.g., an axis) of the variable area inlet assembly  64 , which centerline  82  may be coaxial with the axial centerline  30 . The inner barrel  76  extends axially along the axial centerline  30 ,  82  between a first (e.g., forward, upstream) end  84  of the inner barrel  76  and a second (e.g., aft, downstream) end  86  of the inner barrel  76 . The inner barrel second end  86  of  FIG.  1    is connected to a (e.g., forward, upstream) end of the outer case  38 . The inner barrel  76  may be configured to attenuate noise generated during aircraft propulsion system operation and, more particularly for example, noise generated by rotation of the fan rotor  48 . The inner barrel  76 , for example, may include at least one tubular noise attenuating acoustic panel or a circumferential array of arcuate noise attenuating acoustic panels arranged around the axial centerline. The present disclosure, however, is not limited to such an acoustic inner barrel configuration. 
     The outer barrel  78  extends circumferentially about (e.g., completely around) the axial centerline  30 ,  82 . The outer barrel  78  extends axially along the axial centerline  30 ,  82  between a first (e.g., forward, upstream) end  88  of the outer barrel  78  and a second (e.g., aft, downstream) end  90  of the outer barrel  78 . The outer barrel second end  90  of  FIG.  1    is disposed next to respective (e.g., forward, upstream) ends of a pair of fan cowls (one visible in  FIG.  1   ) of the outer nacelle structure  40 . 
     Referring to  FIG.  5   , the inlet lip  80  forms a leading edge  92  of the nacelle  24  as well as an outer peripheral boundary of the inlet orifice  70  and a forward, upstream portion of the inlet passage  68 . A flat plane  94  defined by at least three points along (or an entirety of) the nacelle leading edge  92  (see plane of  FIG.  2   ) is angularly offset from the axial centerline  30 ,  82  by an angle  96 ; e.g., an acute angle. With this arrangement, a first point  98  (e.g., one of the at least three points) on the nacelle leading edge  92  is axially displaced from a second point  100  (e.g., another one of the at least three points) on the nacelle leading edge  92  by a non-zero axial distance  102 . The first point  98  of  FIG.  5    may be a forwardmost, upstream-most point along the nacelle leading edge  92  located at, for example, a gravitational top of the aircraft propulsion system  20 . The second point  100  of  FIG.  5    may be an aftmost, downstream-most point along the nacelle leading edge  92  and/or diametrically opposed to the first point  98  located at, for example, a gravitational bottom of the aircraft propulsion system  20 . The present disclosure, however, is not limited to such an exemplary inlet lip configuration. For example, in other embodiments, the first point  98  and/or the second point  100  may be respectively located along opposing sides of the aircraft propulsion system  20 . Furthermore, while the first point  98  may be the only forwardmost, upstream-most point along the nacelle leading edge  92  and/or the second point  100  may be the only aftmost, downstream-most point along the nacelle leading edge  92 , the present disclosure is not limited to such an exemplary configuration. The nacelle leading edge  92 , for example, may alternatively have an undulating (e.g., wavy) geometry. 
     The inlet lip  80  of  FIG.  5    has a cupped (e.g., a generally U-shaped or V-shaped) side sectional geometry. The inlet lip  80  and its cupped side sectional geometry extend circumferentially about (e.g., completely around) the axial centerline  30 ,  82 ; see also  FIG.  2   . The inlet lip  80  of  FIG.  5   , for example, includes axially overlapping inner and outer lip portions  104  and  106 . The inner lip portion  104  is connected to and may be integral with the outer lip portion  106  at and along the nacelle leading edge  92 . An (e.g., aft, downstream) end of the inner lip portion  104  is axially adjacent and/or connected to the inner barrel first end  84 . An (e.g., aft, downstream) end of the outer lip portion  106  is axially adjacent and/or connected to the outer barrel first end  88 . 
     Referring to  FIG.  6   , the center body  66  may be configured as a double tapered center body. The center body  66  of  FIG.  6   , for example, has a double tapered exterior skin that extends circumferentially about (e.g., completely around) an axial centerline  108  (e.g., a rotational axis) of the center body  66 , which centerline  108  may be coaxial with the axial centerline  30 ,  82  (see also  FIG.  2   ). The center body  66  and its exterior surface extend axially along the axial centerline  30 ,  82 ,  108  from a first (e.g., forward, upstream) end  110  of the center body  66  to a second (e.g., aft, downstream) end  112  of the center body  66 . The exterior skin of  FIG.  6    includes a first (e.g., forward, upstream) surface  114 , a second (e.g., aft, downstream) surface  116  and a (e.g., intermediate) plateau surface  118 . 
     The first surface  114  may be a tapered and/or tubular surface. The first surface  114  of  FIG.  6   , for example, is configured as a substantially conical surface. This first surface  114  extends circumferentially about (e.g., completely around) the axial centerline  30 ,  82 ,  108 . The first surface  114  extends axially along the axial centerline  30 ,  82 ,  108  from a leading end tip (e.g., forward, upstream point) of the center body  66  at the center body first end  110  to a (e.g., annular) leading edge  120  of the plateau surface  118 . The first surface  114  flares (e.g., continuously or intermittently) radially away from the axial centerline  30 ,  82 ,  108  as the center body  66  and its surface  114  extend axially in an aft, downstream direction along the axial centerline  30 ,  82 ,  108  from the leading end tip towards (e.g., to) the plateau surface  118 . The first surface  114  thereby tapers (e.g., continuously or intermittently) radially towards the axial centerline  30 ,  82 ,  108  as the center body  66  and its surface  114  extend axially in a forward, upstream direction along the axial centerline  30 ,  82 ,  108  from the plateau surface  118  towards (e.g., to) the leading end tip. 
     The second surface  116  may be a tapered and/or tubular surface. The second surface  116  of  FIG.  6   , for example, is configured as a substantially conical surface. This second surface  116  extends circumferentially about (e.g., completely around) the axial centerline  30 ,  82 ,  108 . The second surface  116  extends axially along the axial centerline  30 ,  82 ,  108  from a trailing edge of the center body  66  at the center body second end  112  to a (e.g., annular) trailing edge  122  of the plateau surface  118 . The second surface  116  tapers (e.g., continuously or intermittently) radially towards the axial centerline  30 ,  82 ,  108  as the center body  66  and its surface  116  extend axially in the aft, downstream direction along the axial centerline  30 ,  82 ,  108  from the plateau surface  118  towards (e.g., to) the center body trailing edge. The second surface  116  thereby flares (e.g., continuously or intermittently) radially away from the axial centerline  30 ,  82 ,  108  as the center body  66  and its surface  116  extend axially in the forward, upstream direction along the axial centerline  30 ,  82 ,  108  from the center body trailing edge towards (e.g., to) the plateau surface  118 . 
     The plateau surface  118  may be an apex surface of the center body  66 . The plateau surface  118  of  FIG.  6   , for example, is configured as a substantially cylindrical surface. This plateau surface  118  extends circumferentially about (e.g., completely around) the axial centerline  30 ,  82 ,  108 . The plateau surface  118  extends axially along the axial centerline  30 ,  82 ,  108  from the plateau surface leading edge  120  to the plateau surface trailing edge  122 . A flat plane  124  defined by at least three points along (or an entirety of) the plateau surface leading edge  120  is angularly offset from the axial centerline  30 ,  82 ,  108  by a first angle  126 ; e.g., a right angle. A flat plane  128  defined by at least three points along (or an entirety of) the plateau surface trailing edge  122  is angularly offset from the axial centerline  30 ,  82 ,  108  by a second angle  130 ; e.g., an acute angle. The second angle  130  of  FIG.  6    is different (e.g., less) than the first angle  126 . The plateau surface trailing edge plane  128  of  FIG.  6    is thereby angularly offset from the plateau surface leading edge plane  124 . The plateau surface trailing edge plane  128  may (or may not) be parallel with the inlet structure leading edge plane  94  (see  FIG.  1   ). Thus, the plateau surface trailing edge plane  128  may have a scarfed configuration. 
     A first point  132  (e.g., one of the at least three points) on the plateau surface trailing edge  122  is axially displaced from a second point  134  (e.g., another one of the at least three points) on the plateau surface trailing edge  122  by a non-zero axial distance  136 . The first point  132  of  FIG.  6    may be a forwardmost, upstream-most point along the trailing edge  122  located at, for example, the gravitational top of the aircraft propulsion system  20  during the first (e.g., supersonic) mode of operation. The second point  134  of  FIG.  6    may be an aftmost, downstream-most point along the trailing edge  122  and/or diametrically opposed to the first point  132  located at, for example, the gravitational bottom of the aircraft propulsion system  20  during the first (e.g., supersonic) mode of operation. The present disclosure, however, is not limited to such an exemplary inlet lip configuration. For example, in other embodiments, the first point  132  and/or the second point  134  may be respectively located along opposing sides of the center body  66  during the first (e.g., supersonic) mode of operation. Furthermore, while the first point  132  may be the only forwardmost, upstream-most point along the plateau surface trailing edge  122  and/or the second point  134  may be the only aftmost, downstream-most point along the plateau surface trailing edge  122 , the present disclosure is not limited to such an exemplary configuration. The plateau surface leading edge  120 , for example, may alternatively have an undulating (e.g., wavy) geometry. 
     The plateau surface  118  has an axial length  138 . This plateau surface axial length  138  (e.g., continuously or intermittently) changes as the plateau surface  118  extends circumferentially about (e.g., completely around) the axial centerline  30 ,  82 ,  108 . The plateau surface axial length  138  may have a first (e.g., maximum) value at the first point  132 . The plateau surface axial length  138  may have a second (e.g., minimum) value at the second point  134 , which second value is different (e.g., less) than the first value. 
     The center body  66  of  FIGS.  1  and  2    is arranged within the nacelle inlet structure  58  and projects axially through the inlet orifice  70 . The center body  66  at least partially (or completely) forms an inner peripheral boundary of the inlet passage  68 . The plateau surface  118  is axially aligned with the inlet lip  80 . 
     The center body  66  may be configured to rotate (e.g., clockwise or counter-clockwise) about the axial centerline  30 ,  82 ,  108  between the first position (see  FIG.  7   ) and the second position (see  FIG.  8   ). The center body  66 , for example, may be rotationally supported by one or more bearing structures, track assemblies and/or other suitable slidable/movable/pivotable connectors (not shown). 
     In the first position of  FIG.  7   , an inner peripheral boundary of the metering portion  72  is (e.g., completely) formed by the plateau surface  118 . In this position, a minimum distance  140  between the inlet lip  80  and the center body  66  and, more particularly, between the inner lip portion  104  and the plateau surface  118  may be at least substantially or completely equal/uniform about the axial centerline  30 ,  82 ,  108 . An axial distance  142  between the nacelle leading edge  92  and the plateau surface trailing edge  122  may also or alternatively be at least substantially or completely equal/uniform about the axial centerline  30 ,  82 ,  108 . 
     In the second position of  FIG.  8   , the inner peripheral boundary of the metering portion  72  is formed by at least (or only) a portion of the plateau surface  118  and a portion of the second surface  116 . In this position, the minimum distance  140  may (e.g., continuously or intermittently) change about the axial centerline  30 ,  82 ,  108 . For example, the minimum distance  140  at a first point  144  along the metering portion  72  may be different (e.g., less) than the minimum distance at a second point  146  along the metering portion  72 . The first point  144  of  FIG.  8    may be at the gravitational top of the aircraft propulsion system  20 . The second point  146  of  FIG.  8    may be diametrically opposed to the first point  144  and/or at the gravitational bottom of the aircraft propulsion system  20 . The minimum distance  140  at the first point  144  may be equal to the minimum distance  140  when the center body  66  is in the first position of  FIG.  7   . The axial distance  142  between the nacelle leading edge  92  and the plateau surface trailing edge  122  may also or alternatively (e.g., continuously or intermittently) change about the axial centerline  30 ,  82 ,  108 . For example, the axial distance  142  at the first point  144  may be different (e.g., greater) than the axial distance  142  at the second point  146 . With the foregoing arrangement, the second area  74 B of the metering portion  72  (see  FIG.  4   ) is different (e.g., greater) than the first area  74 A of the metering portion  72  (see  FIG.  3   ). 
     In some embodiments, the axial centerline/rotational axis  108  of the center body  66  is coaxial with the axial centerline/rotational axis  30  of the aircraft propulsion system  20  and its gas turbine engine  22  as described above. However, in other embodiments, the axial centerline/rotational axis  108  of the center body  66  may be eccentric/non-coaxial with the axial centerline/rotational axis  30  of the aircraft propulsion system  20  and its gas turbine engine  22  as shown in  FIG.  9   . More particularly, the axial centerline  108  may be displaced from and/or angularly offset from the axial centerline  30 . 
     In some embodiments, referring to  FIGS.  7  and  8   , the center body  66  may rotate one-hundred and eighty degrees (180°) between the first position of  FIG.  7    and the second position of  FIG.  8   . However, in other embodiments, the center body  66  may rotate more than hundred and eighty degrees (e.g., between hundred and eighty degrees (180°) and two-hundred and seventy degrees (270°)) between the first and the second positions. In still other embodiments, the center body  66  may rotate less than hundred and eighty degrees (e.g., between ninety degrees (90°) and hundred and eighty degrees (180°)) between the first and the second positions. 
     In some embodiments, the center body  66  may be actuated by a gear drive system. In other embodiments, the center body  66  may be actuated by one or more other types of actuators such as, but not limited to, one or more worm and gear arrangements and/or one or more linear actuators arranged around a periphery of the center body  66 . 
     The center body  66  is described above as a movable structure, and the nacelle inlet structure  58  is described above as a stationary structure. However, it is contemplated that the functionality/operation of these structures  58  and  66  may be reversed. For example, in some embodiments, the center body  66  may be configured as a stationary structure, and the nacelle inlet structure  58  and/or its inlet lip  80  may be configured as a movable structure; e.g., a rotatable structure. The nacelle inlet structure  58  and/or its inlet lip  80  may thereby move (e.g., rotate clockwise or counter-clockwise about the centerline  30 ,  82 ,  108 ) between the first and the second positions to increase or decrease the area of the metering portion  72 . 
     The aircraft propulsion system  20  and its variable area inlet assembly  64  may be configured with various gas turbine engines other than the one described above. The gas turbine engine  22 , for example, may be configured as a geared or a direct drive turbine engine. The gas turbine engine  22  may be configured with a single spool, with two spools (e.g., see  FIG.  1   ), or with more than two spools. The gas turbine engine  22  may be configured as a turbofan engine, a turbojet engine or any other type of turbine engine. The present invention therefore is not limited to any particular types or configurations of gas turbine engines. The present disclosure is also not limited to applications where the aircraft is capable to traveling supersonic speeds. 
     While various embodiments of the present invention have been disclosed, it will be apparent to those of ordinary skill in the art that many more embodiments and implementations are possible within the scope of the invention. For example, the present invention as described herein includes several aspects and embodiments that include particular features. Although these features may be described individually, it is within the scope of the present invention that some or all of these features may be combined with any one of the aspects and remain within the scope of the invention. Accordingly, the present invention is not to be restricted except in light of the attached claims and their equivalents.