Patent Description:
A conventional commercial aircraft generally includes a fuselage, a pair of wings, and a propulsion system that provides thrust. The propulsion system typically includes at least two aircraft engines, such as turbofan jet engines. Each turbofan jet engine is mounted to a respective one of the wings of the aircraft, such as in a suspended position beneath the wing, separated from the wing and fuselage. Such a configuration allows for the turbofan jet engines to interact with separate, freestream airflows that are not impacted by the wings and/or fuselage. This configuration can reduce an amount of turbulence within the air entering an inlet of each respective turbofan jet engine, which has a positive effect on a net propulsive thrust of the aircraft.

However, a drag on the aircraft including the turbofan jet engines also affects the net propulsive thrust of the aircraft. A total amount of drag on the aircraft, including skin friction and form drag, is generally proportional to a difference between a freestream velocity of air approaching the aircraft and an average velocity of a wake downstream from the aircraft that is produced due to the drag on the aircraft.

Positioning a fan at an aft end of the fuselage of the aircraft may assist with reenergizing a boundary layer airflow over the aft end of the fuselage and improving propulsive efficiency. However, given existing structures at the aft end of the fuselage, such as one or more stabilizers, the airflow ingested by such a fan may not have a uniform velocity or total pressure profile along the circumferential and radial directions of the fan. More specifically, the structures at the aft end of the fuselage may generate a boundary layer or wake resulting in swirl distortion and a distorted velocity or total pressure profile of the airflow ingested by the fan.

Accordingly, an aircraft capable of energizing slow-moving air forming a boundary layer across the fuselage of the aircraft would be useful. Specifically, a fuselage of an aircraft designed to increase the ingestion of relatively low momentum boundary layer airflow into the aft engine and reduce the non-uniformity and distortion of the velocity profile of ingested airflow would be particularly beneficial. <CIT> describes an aircraft having an aft-mounted fan for ingesting boundary layer air.

Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures, <FIG> provides a top view of an exemplary aircraft <NUM> as may incorporate various embodiments of the present invention. <FIG> provides a port side <NUM> view of the aircraft <NUM> as illustrated in <FIG>. As shown in <FIG> collectively, the aircraft <NUM> defines a longitudinal direction <NUM> that extends therethrough, a vertical direction V, a lateral direction L, a forward end <NUM>, and an aft end <NUM>. Moreover, the aircraft <NUM> defines a mean line <NUM> extending between the forward end <NUM> and aft end <NUM> of the aircraft <NUM>. As used herein, the "mean line" refers to a midpoint line extending along a length of the aircraft <NUM>, not taking into account the appendages of the aircraft <NUM> (such as the wings <NUM> and stabilizers discussed below).

Moreover, the aircraft <NUM> includes a fuselage <NUM>, extending longitudinally from the forward end <NUM> of the aircraft <NUM> towards the aft end <NUM> of the aircraft <NUM>, and a pair of wings <NUM>. As used herein, the term "fuselage" generally includes all of the body of the aircraft <NUM>, such as an empennage of the aircraft <NUM> and an outer surface or skin of the aircraft <NUM>. The first of such wings <NUM> extends laterally outwardly with respect to the longitudinal direction <NUM> from the port side <NUM> of the fuselage <NUM> and the second of such wings <NUM> extends laterally outwardly with respect to the longitudinal direction <NUM> from a starboard side <NUM> of the fuselage <NUM>. Each of the wings <NUM> for the exemplary embodiment depicted includes one or more leading edge flaps <NUM> and one or more trailing edge flaps <NUM>. The aircraft <NUM> further includes a vertical stabilizer <NUM> having a rudder flap <NUM> for yaw control, and a pair of horizontal stabilizers <NUM>, each having an elevator flap <NUM> for pitch control. The fuselage <NUM> additionally includes an outer surface <NUM>.

As illustrated, each stabilizer extends between a root portion and a tip portion substantially within a single plane. For example, as illustrated in <FIG>, vertical stabilizer <NUM> defines a root portion <NUM> and a tip portion <NUM> separated along the vertical direction V. In addition, vertical stabilizer <NUM> extends between a leading edge <NUM> and a trailing edge <NUM> along the longitudinal direction <NUM>. As illustrated, vertical stabilizer <NUM> is mounted to fuselage <NUM> at root portion <NUM> and extends substantially along the vertical direction V to tip portion <NUM>. In this manner, a junction line <NUM> is defined at the intersection of vertical stabilizer <NUM> and fuselage <NUM>. More specifically, junction line <NUM> extends between leading edge <NUM> and trailing edge <NUM> of vertical stabilizer <NUM>. However, it should be appreciated that in other exemplary embodiments of the present disclosure, the aircraft <NUM> may additionally or alternatively include any other suitable configuration of stabilizers that may or may not extend directly along the vertical direction V or horizontal/lateral direction L. In addition, alternative stabilizers may be any suitable shape, size, configuration, or orientation while remaining within the scope of the present subject matter.

The exemplary aircraft <NUM> of <FIG> also includes a propulsion system. The exemplary propulsion system includes a plurality of aircraft engines, at least one of which mounted to each of the pair of wings <NUM>. Specifically, the plurality of aircraft engines includes a first aircraft engine <NUM> mounted to a first wing of the pair of wings <NUM> and a second aircraft engine <NUM> mounted to a second wing of the pair of wings <NUM>. In at least certain exemplary embodiments, the aircraft engines <NUM>, <NUM> may be configured as turbofan jet engines suspended beneath the wings <NUM> in an under-wing configuration. For example, in at least certain exemplary embodiments, the first and/or second aircraft engines <NUM>, <NUM> may be configured in substantially the same manner as the exemplary turbofan jet engine <NUM> described below with reference to <FIG>. Alternatively, however, in other exemplary embodiments any other suitable aircraft engine may be provided. For example, in other exemplary embodiments the first and/or second aircraft engines <NUM>, <NUM> may alternatively be configured as turbojet engines, turboshaft engines, turboprop engines, etc..

Additionally, the propulsion system includes an aft engine <NUM> mounted to the fuselage <NUM> of the aircraft <NUM> at the aft end <NUM> of the fuselage <NUM>. The exemplary aft engine <NUM> is mounted to the fuselage <NUM> of the aircraft <NUM> such that the mean line <NUM> extends therethrough. The aft engine <NUM>, which is generally configured as an engine that ingests and consumes air forming a boundary layer over fuselage <NUM>, will be discussed in greater detail below with reference to <FIG>.

Referring specifically to <FIG>, the aircraft <NUM> additionally includes landing gear, such as wheels <NUM>, extending from a bottom side of the fuselage <NUM> and from a bottom side of the wings <NUM>. The fuselage <NUM> is designed to allow the aircraft <NUM> to takeoff and/or land at a takeoff angle <NUM> with the ground without the aft end <NUM> scraping the ground. More specifically, takeoff angle <NUM> may be defined as the angle between the ground (parallel to longitudinal direction <NUM>) and a takeoff plane <NUM>. As will be discussed below, the exemplary fuselage <NUM> and aft engine <NUM> described herein are designed to allow the aircraft <NUM> to maintain a desired takeoff angle <NUM>, despite the addition of the aft engine <NUM> proximate the aft end <NUM> of the aircraft <NUM>. Notably, for the embodiment depicted, the longitudinal direction <NUM> of the aircraft <NUM> is parallel to the ground when the aircraft <NUM> is on the ground. Accordingly, the maximum takeoff angle <NUM>, as shown, may alternatively be defined with the longitudinal direction <NUM> of the aircraft <NUM> (shown as angle <NUM>' in <FIG>).

Referring now to <FIG>, a schematic, cross-sectional view of an exemplary aircraft engine is provided. Specifically, for the embodiment depicted, the aircraft engine is configured as a high bypass turbofan jet engine, referred to herein as "turbofan engine <NUM>. " As discussed above, one or both of the first and/or second aircraft engines <NUM>, <NUM> of the exemplary aircraft <NUM> described in <FIG> may be configured in substantially the same manner as the exemplary turbofan engine <NUM> of <FIG>. Alternatively, however, in other exemplary embodiments, one or both of aircraft engines <NUM>, <NUM> may be configured as any other suitable engines, such as a turboshaft, turboprop, turbojet, etc..

As shown in <FIG>, the turbofan engine <NUM> defines an axial direction A<NUM> (extending parallel to a longitudinal centerline <NUM> provided for reference) and a radial direction R<NUM>. In general, the turbofan <NUM> includes a fan section <NUM> and a core turbine engine <NUM> disposed downstream from the fan section <NUM>.

The exemplary core turbine engine <NUM> depicted generally includes a substantially tubular outer casing <NUM> that defines an annular inlet <NUM>. The outer casing <NUM> encases, in serial flow relationship, a compressor section including a booster or low pressure (LP) compressor <NUM> and a high pressure (HP) compressor <NUM>; a combustion section <NUM>; a turbine section including a high pressure (HP) turbine <NUM> and a low pressure (LP) turbine <NUM>; and a jet exhaust nozzle section <NUM>. A high pressure (HP) shaft or spool <NUM> drivingly connects the HP turbine <NUM> to the HP compressor <NUM>. A low pressure (LP) shaft or spool <NUM> drivingly connects the LP turbine <NUM> to the LP compressor <NUM>. The compressor section, combustion section <NUM>, turbine section, and nozzle section <NUM> together define a core air flowpath.

For the embodiment depicted, the fan section <NUM> includes a variable pitch fan <NUM> having a plurality of fan blades <NUM> coupled to a disk <NUM> in a spaced apart manner. As depicted, the fan blades <NUM> extend outwardly from disk <NUM> generally along the radial direction R<NUM> and define a fan diameter D. Each fan blade <NUM> is rotatable relative to the disk <NUM> about a pitch axis P by virtue of the fan blades <NUM> being operatively coupled to a suitable actuation member <NUM> configured to collectively vary the pitch of the fan blades <NUM> in unison. The fan blades <NUM>, disk <NUM>, and actuation member <NUM> are together rotatable about the longitudinal direction <NUM> by LP shaft <NUM> across a power gear box <NUM>. The power gear box <NUM> includes a plurality of gears for adjusting the rotational speed of the fan <NUM> relative to the LP shaft <NUM> to a more efficient rotational fan speed.

Referring still to the exemplary embodiment of <FIG>, the disk <NUM> is covered by rotatable front hub <NUM> aerodynamically contoured to promote an airflow through the plurality of fan blades <NUM>. Additionally, the exemplary fan section <NUM> includes an annular fan casing or outer nacelle <NUM> that circumferentially surrounds the fan <NUM> and/or at least a portion of the core turbine engine <NUM>. It should be appreciated that the nacelle <NUM> may be configured to be supported relative to the core turbine engine <NUM> by a plurality of circumferentially-spaced outlet guide vanes <NUM>. Moreover, a downstream section <NUM> of the nacelle <NUM> may extend over an outer portion of the core turbine engine <NUM> so as to define a bypass airflow passage <NUM> therebetween.

It should be appreciated, however, that the exemplary turbofan engine <NUM> depicted in <FIG> is by way of example only, and that in other exemplary embodiments, the turbofan engine <NUM> may have any other suitable configuration, including, e.g., any suitable number of shafts or spools, compressors, and/or turbines.

Referring now also to <FIG>, a close-up, schematic, cross-sectional view of the exemplary aft engine <NUM> of <FIG> is provided. As discussed, the exemplary aft engine <NUM> is mounted to the fuselage <NUM> proximate the aft end <NUM> of the aircraft <NUM>. The aft engine <NUM> depicted defines an axial direction A<NUM> extending along a longitudinal centerline axis <NUM> that extends therethrough for reference, a radial direction R<NUM>, and a circumferential direction C<NUM> (see <FIG>).

Additionally, for the embodiment depicted, the aft engine <NUM> is configured as a boundary layer ingestion engine configured to ingest and consume air forming a boundary layer over the fuselage <NUM> of the aircraft <NUM>. The aft engine <NUM> includes a fan <NUM> rotatable about the centerline axis <NUM>, a nacelle <NUM> extending around a portion of the fan <NUM>, and one or more structural members <NUM> extending between the nacelle <NUM> and the fuselage <NUM> of the aircraft <NUM>. The fan <NUM> includes a plurality of fan blades <NUM> spaced generally along circumferential direction C<NUM>. Additionally, the nacelle <NUM> extends around and encircles the plurality of fan blades <NUM> and a portion of the fuselage <NUM>. Specifically, the nacelle <NUM> extends around at least a portion of the fuselage <NUM> of the aircraft <NUM> when, as in <FIG>, the aft engine <NUM> is mounted to the aircraft <NUM>.

As is also depicted in <FIG>, the fan <NUM> further includes a fan shaft <NUM> with the plurality of fan blades <NUM> attached thereto. Although not depicted, the fan shaft <NUM> may be rotatably supported by one or more bearings located forward of the plurality of fan blades <NUM> and, optionally, one or more bearings located aft of the plurality of fan blades <NUM>. Such bearings may be any suitable combination of roller bearings, ball bearings, thrust bearings, etc..

In certain exemplary embodiments, the plurality of fan blades <NUM> may be attached in a fixed manner to the fan shaft <NUM>, or alternatively, the plurality of fan blades <NUM> may be rotatably attached to the fan shaft <NUM>. For example, the plurality of fan blades <NUM> may be attached to the fan shaft <NUM> such that a pitch of each of the plurality of fan blades <NUM> may be changed, e.g., in unison, by a pitch change mechanism (not shown).

The fan shaft <NUM> is mechanically coupled to a power source <NUM> located at least partially within the fuselage <NUM> of the aircraft <NUM>. For the embodiment depicted, the fan shaft <NUM> is mechanically coupled to the power source <NUM> through a gearbox <NUM>. The gearbox <NUM> may be configured to modify a rotational speed of the power source <NUM>, or rather of a shaft <NUM> of the power source <NUM>, such that the fan <NUM> of the aft engine <NUM> rotates at a desired rotational speed. The gearbox <NUM> may be a fixed ratio gearbox, or alternatively, the gearbox <NUM> may define a variable gear ratio.

The power source <NUM> may be any suitable power source. For example, in certain exemplary embodiments the power source <NUM> may be an electric power source (e.g., the aft engine <NUM> may be configured as part of a gas-electric propulsion system with the first and/or second aircraft engines <NUM>, <NUM>). However, in other exemplary embodiments, the power source <NUM> may alternatively be configured as a dedicated gas engine, such as a gas turbine engine. Moreover, in certain exemplary embodiments, the power source <NUM> may be positioned at any other suitable location within, e.g., the fuselage <NUM> of the aircraft <NUM> or the aft engine <NUM>. For example, in certain exemplary embodiments, the power source <NUM> may be configured as a gas turbine engine positioned at least partially within the aft engine <NUM>.

Referring still to <FIG>, the one or more structural members <NUM> extend between the nacelle <NUM> and the fuselage <NUM> of the aircraft <NUM> at a location forward of the plurality of fan blades <NUM>. The one or more structural members <NUM> for the embodiment depicted extend substantially along the radial direction R<NUM> between the nacelle <NUM> and the fuselage <NUM> of the aircraft <NUM> for mounting the aft engine <NUM> to the fuselage <NUM> of the aircraft <NUM>. It should also be appreciated, however, that in other exemplary embodiments the one or more structural members <NUM> may instead extend substantially along the axial direction A<NUM>, or in any other suitable direction between the axial and radial directions A<NUM>, R<NUM>. It should be appreciated, that as used herein, terms of approximation, such as "approximately," "substantially," or "about," refer to being within a ten percent margin of error.

The one or more structural members <NUM> depicted are configured as inlet guide vanes for the fan <NUM>, such that the one or more structural members <NUM> are shaped and oriented to direct and condition a flow of air into the aft engine <NUM> to increase an efficiency of the aft engine <NUM>. In certain exemplary embodiments, the one or more structural members <NUM> may be configured as fixed inlet guide vanes extending between the nacelle <NUM> and the fuselage <NUM> of the aircraft <NUM>, or alternatively the one or more structural members <NUM> may be configured as variable inlet guide vanes.

Moreover, the aft engine <NUM> includes one or more outlet guide vanes <NUM> and a tail cone <NUM>. The one or more outlet guide vanes <NUM> for the embodiment depicted extend between the nacelle <NUM> and the tail cone <NUM> for, e.g., adding strength and rigidity to the aft engine <NUM>, as well as for directing a flow of air through the aft engine <NUM>. The outlet guide vanes <NUM> may be evenly spaced along the circumferential direction C<NUM> (see <FIG>), or may have any other suitable spacing. Additionally, the outlet guide vanes <NUM> may be fixed outlet guide vanes, or alternatively may be variable outlet guide vanes.

Aft of the plurality of fan blades <NUM>, and for the embodiment depicted, aft of the one or more outlet guide vanes <NUM>, the aft engine <NUM> additionally defines a nozzle <NUM> between the nacelle <NUM> and the tail cone <NUM>. The nozzle <NUM> may be configured to generate an amount of thrust from the air flowing therethrough, and the tail cone <NUM> may be shaped to minimize an amount of drag on the aft engine <NUM>. However, in other embodiments, the tail cone <NUM> may have any other shape and may, e.g., end forward of an aft end of the nacelle <NUM> such that the tail cone <NUM> is enclosed by the nacelle <NUM> at an aft end. Additionally, in other embodiments, the aft engine <NUM> may not be configured to generate any measureable amount of thrust, and instead may be configured to ingest air from a boundary layer of air of the fuselage <NUM> of the aircraft <NUM> and add energy/ speed up such air to reduce an overall drag on the aircraft <NUM> (and thus increase a net thrust of the aircraft <NUM>).

Referring still to <FIG>, the aft engine <NUM>, or rather the nacelle <NUM>, defines an inlet <NUM> at a forward end <NUM> of the nacelle <NUM>. The inlet <NUM> is defined by the nacelle <NUM> with the fuselage <NUM>, i.e., between the nacelle <NUM> and the fuselage <NUM>. As mentioned above, the nacelle <NUM> of the aft engine <NUM> extends around and surrounds the plurality of fan blades <NUM> of the fan <NUM> of the aft engine <NUM>. For the embodiment depicted, nacelle <NUM> also extends at least partially around the central axis <NUM> of the aft engine <NUM>, and at least partially around the mean line <NUM> of the aircraft <NUM>. Specifically, for the embodiment depicted, the nacelle <NUM> extends substantially three hundred and sixty degrees (<NUM>°) around the central axis <NUM> of the aft engine <NUM>, and substantially three hundred and sixty degrees (<NUM>°) around the mean line <NUM> of the aircraft <NUM>.

Notably, by positioning the aft engine <NUM> such that the nacelle <NUM> of the aft engine <NUM> extends at least partially around the fuselage <NUM> proximate the aft end <NUM> of the aircraft <NUM>, a bottom portion <NUM> of the nacelle <NUM> may not interfere with, e.g., the takeoff angle <NUM> of the aircraft <NUM> (see <FIG>). For example, as shown, the nacelle <NUM> of the aft engine <NUM> includes at least a portion located inward of the takeoff plane <NUM> defined by the fuselage <NUM> (see <FIG>). Particularly for the embodiment depicted, an entirety of the bottom portion <NUM> of the nacelle <NUM> is positioned in-line with, or inwardly of the takeoff plane <NUM> of the fuselage <NUM>. For at least certain prior art aircrafts, the takeoff plane <NUM> of the fuselage <NUM> indicates the conventional shape for a bottom portion of a fuselage at an aft end of an aircraft.

Referring now to <FIG>, the shape of the aft end <NUM> of the exemplary aircraft <NUM> as well as features for providing improved boundary layer ingestion will be described in more detail. More specifically, <FIG> and <FIG> provide schematic, cross-sectional side views of aft engine <NUM> mounted to fuselage <NUM>. <FIG> provide schematic cross-sectional views of fuselage, taken along the Line X-X in <FIG>.

Referring specifically to <FIG>, according to an exemplary embodiment, top side <NUM> of fuselage <NUM> defines a top surface <NUM> along which boundary layer air flows over aircraft <NUM>. Similarly, bottom side <NUM> defines a bottom surface <NUM> along which boundary layer air flows over aircraft <NUM>. As explained above, it is desirable to accelerate low velocity boundary layer airflow to improve propulsive efficiency. The features of the aircraft <NUM> described herein achieve these and other objectives.

According to the illustrated embodiment, top surface <NUM> defines a first point <NUM> located in a plane perpendicular to the longitudinal direction <NUM> and positioned at or aft of where leading edge <NUM> of vertical stabilizer <NUM> meets fuselage <NUM>. In addition, top surface <NUM> defines a second point <NUM> located in a plane perpendicular to the longitudinal direction <NUM> downstream of first point <NUM>. For example, second point <NUM> may be positioned at or forward of where trailing edge <NUM> of vertical stabilizer <NUM> meets fuselage <NUM>. Top surface <NUM> also defines an upper inflection point <NUM> positioned between first point <NUM> and second point <NUM> along top surface <NUM> of fuselage <NUM>. As illustrated, a first portion <NUM> of top surface <NUM> extends between first point <NUM> and upper inflection point <NUM> and a second portion <NUM> of top surface <NUM> extends between upper inflection point <NUM> and second point <NUM>.

As illustrated in <FIG>, first portion <NUM> is a convex curve when viewed looking down onto top surface <NUM> from outside of fuselage <NUM>. In addition, second portion <NUM> is a concave curve when viewed looking down onto top surface <NUM> from outside of fuselage <NUM>. In this regard, fuselage <NUM> generally defines a convex surface upstream a concave surface proximate aft engine <NUM>. In this manner, boundary layer airflow may more effectively be distributed within aft engine <NUM>. Thus, aft engine <NUM> may ingest a desired amount of slower moving boundary layer airflow and may discharge that low velocity air as relatively higher velocity air, thereby improving the propulsive efficiency of aircraft <NUM>.

It should be appreciated that bottom surface <NUM> and any other surface located circumferentially around fuselage <NUM> proximate aft end <NUM> of fuselage <NUM> may have a similar profile as top surface <NUM>. For example, bottom surface <NUM> defines a first point <NUM> located, for example, in the same plane as first point <NUM>. In addition, bottom surface <NUM> defines a second point <NUM> located, for example, in the same plane as second point <NUM>. It should be appreciated that first point <NUM> and second point <NUM> may alternatively be positioned at any suitable location along bottom surface <NUM> of fuselage <NUM>. Bottom surface <NUM> also defines a lower inflection point <NUM> positioned between first point <NUM> and second point <NUM> along bottom surface <NUM> of fuselage <NUM>. As illustrated, a convex first portion <NUM> of bottom surface <NUM> extends between first point <NUM> and lower inflection point <NUM> and a concave second portion <NUM> of bottom surface <NUM> extends between lower inflection point <NUM> and second point <NUM>.

As illustrated, fuselage <NUM> defines upper inflection point <NUM> and lower inflection point <NUM> upstream of inlet <NUM> to aft engine <NUM>. According to the illustrated embodiment, upper inflection point <NUM> and lower inflection point <NUM> are defined in the same plane between leading edge <NUM> of vertical stabilizer <NUM> and trailing edge <NUM> of vertical stabilizer <NUM>. For example, upper inflection point <NUM> and lower inflection point <NUM> may be defined at a halfway point between leading edge <NUM> and trailing edge <NUM> of vertical stabilizer <NUM>. However, it should be appreciated that upper inflection point <NUM> and lower inflection point <NUM> may be defined at any suitable location on fuselage <NUM>. For example, upper inflection point <NUM> and lower inflection point <NUM> may be defined in a plane perpendicular to the longitudinal direction <NUM> that is three-quarters of the way along junction line <NUM> from leading edge <NUM> to trailing edge <NUM>. In addition, upper inflection point <NUM> and lower inflection point <NUM> may be positioned at different locations along the longitudinal direction <NUM> (i.e., may be in different vertical planes). It should also be appreciated that the locations of upper inflection point <NUM> and lower inflection point <NUM> discussed herein are used only for explaining aspects of the present subject matter. Other locations and configurations of top surface <NUM> and bottom surface <NUM> of fuselage <NUM> are possible.

Referring still to <FIG>, according to an exemplary embodiment, top surface <NUM> of fuselage <NUM> defines a tangent line <NUM> that extends parallel to top surface <NUM> and intersects a forward lip <NUM> of nacelle <NUM>. According to the illustrated embodiment, tangent line <NUM> is defined where Line X-X intersects fuselage <NUM> (approximately halfway between leading edge <NUM> and trailing edge <NUM> of vertical stabilizer <NUM>). However, other locations are possible. Notably, according to the illustrated embodiment, top surface <NUM> of fuselage <NUM> further defines a recessed portion <NUM> located at the aft end <NUM> just upstream of aft engine <NUM>. Recessed portion <NUM> is defined where top surface <NUM> is indented inwardly toward fuselage <NUM> (i.e., towards the mean line <NUM> of the aircraft <NUM>). However, because relatively higher velocity boundary layer air cannot track recessed portion <NUM> as easily as lower velocity air, the relatively higher velocity air continues along a trajectory defined by tangent line <NUM>, thereby avoiding ingestion by aft engine <NUM>.

It should be appreciated that the shape of fuselage <NUM> illustrated in <FIG> is only one exemplary fuselage <NUM> shape. Referring to <FIG>, fuselage <NUM> may define several regions along the aft end of fuselage <NUM>, each region being concave, convex, or straight and having varying radii of curvature.

More specifically, according to the illustrated exemplary embodiment, fuselage <NUM> defines a first region <NUM> that extends along junction line <NUM> between leading edge <NUM> and a first point along junction line <NUM>. First region <NUM> is convex, e.g., when viewed looking down onto top surface <NUM> from outside of fuselage <NUM>. In addition, first region <NUM> may have a relatively large radius of curvature, i.e., first radius <NUM>. According to an exemplary embodiment, first region <NUM> may further define an average angle along its length that is approximately ten degrees or less relative to longitudinal direction <NUM>.

Fuselage <NUM> also defines a second region <NUM> that extends along junction line <NUM> between first region <NUM> and a second point along junction line <NUM>. Second region <NUM> is also convex, e.g., when viewed looking down onto top surface <NUM> from outside of fuselage <NUM>. Second region <NUM> may have a radius of curvature, i.e., second radius <NUM>, which is relatively small compared to first radius <NUM>. For example, according to one exemplary embodiment, the ratio of first radius <NUM> to second radius <NUM> may be <NUM>:<NUM>, <NUM>:<NUM>, <NUM>:<NUM>, or greater. Furthermore, according to an exemplary embodiment, second region <NUM> may further define an average angle along its length that is approximately twenty degrees or less relative to longitudinal direction <NUM>.

Fuselage <NUM> also defines a third region <NUM> that extends along junction line <NUM> from second region <NUM> towards end of junction line <NUM>. For example, third region <NUM> may terminate at the end of junction line <NUM>, or at any other location forward of fan <NUM>. Third region <NUM> is concave and may have a radius of curvature, i.e., third radius <NUM>, which is relatively large compared to second radius <NUM>. For example, third radius <NUM> may be approximately the same as first radius <NUM>. It should be appreciated that the regions described above are only used for the purpose of explaining aspects of the present subject matter. There may be fewer or more than three distinct regions, and each may be concave, convex, or have any suitable radius of curvature.

Now referring to <FIG>, two alternative cross-sectional views of fuselage <NUM> will be described according to exemplary embodiments of the present subject matter. Although the profiles used to describe the cross sections of fuselage are different, similar reference numerals will be used to describe them. It should also be appreciated that the cross sections discussed herein are used only for explaining aspects of the present subject matter and are not intended to be limiting in scope. The cross sectional profiles of fuselage <NUM> may vary along the length of fuselage <NUM> as desired depending on the particular application to improve the ingestion of boundary layer airflow into aft engine <NUM>.

Referring now specifically to <FIG>, a first cross section <NUM> will be described according to an exemplary embodiment of the present subject matter. According to the illustrated embodiment, cross section <NUM> may be taken along Line X-X of <FIG>. However, it should be appreciated that cross section <NUM> may be located at any suitable location of fuselage <NUM> along the longitudinal direction <NUM>. For example, cross section <NUM> may be defined at a halfway point between leading edge <NUM> and trailing edge <NUM> of vertical stabilizer <NUM>. Alternatively, cross section <NUM> may be defined at a location along the longitudinal direction <NUM> that is three-quarters of the way along junction line <NUM> from leading edge <NUM> to trailing edge <NUM>.

As illustrated, cross section <NUM> defines a horizontal reference line <NUM> that extends along the lateral direction L between the sides of cross section <NUM>. In addition, horizontal reference line <NUM> extends through the central axis <NUM> of the aft engine <NUM> (see also <FIG>). In this manner, horizontal reference line <NUM> defines a top half <NUM> of cross section <NUM> positioned above horizontal reference line <NUM> along the vertical direction V. In addition, horizontal reference line <NUM> defines a bottom half <NUM> of cross section <NUM> positioned below horizontal reference line <NUM> along the vertical direction V. Notably, according to the illustrated embodiment, a top half cross sectional area of top half <NUM> is greater than a bottom half cross sectional area of bottom half <NUM>. In this manner, inlet <NUM> may be configured to capture a sufficient and uniform amount of the boundary layer air flowing over fuselage <NUM>. For example, according to an exemplary embodiment, the top half cross sectional area of top half <NUM> may be at least about ten percent greater than the bottom half cross sectional area of bottom half <NUM>.

As also illustrated in <FIG>, fuselage <NUM> further defines a reference circle <NUM>. Reference circle <NUM> is defined in the same plane as cross section <NUM>, has a center point <NUM> that corresponds with central axis <NUM>, and has a diameter equivalent to a length of horizontal reference line <NUM> or slightly longer (e.g., less than twenty percent longer) than a length of horizontal reference line <NUM>. Cross section <NUM> defines a circumference <NUM> and reference circle <NUM> defines a circumference <NUM>. According to the illustrated embodiment, at least a portion of the circumference <NUM> of top half <NUM> of cross section <NUM> is located outside reference circle <NUM>. In addition, at least a portion of the circumference <NUM> of bottom half <NUM> of cross section <NUM> is located inside reference circle <NUM>. In this manner, cross section <NUM> may generally be thicker or have a larger cross sectional area on top half <NUM> relative to bottom half <NUM>. Cross section <NUM> may be designed to displace the boundary layer airflow to maximize the ingestion of low velocity air by the aft engine <NUM> and improve the propulsive efficiency of aircraft <NUM>.

As illustrated in <FIG>, cross section <NUM> may be displaced from reference circle <NUM>, such that fuselage <NUM> has an improved profile for boundary layer ingestion. The cross sectional profile may be similar for cross sections taken at other locations along the longitudinal direction <NUM> or may vary depending on the application. However, according to an exemplary embodiment, circumference <NUM> of cross section <NUM> is equivalent to circumference <NUM> of reference circle <NUM>. In this manner, the surface drag along a fuselage shaped as cross section <NUM> may be substantially similar to the surface drag along a fuselage shaped as reference circle <NUM>.

According to an alternative embodiment, horizontal reference line <NUM> extends across a widest portion of cross section <NUM> along the lateral direction L. In such an embodiment, horizontal reference line <NUM> may or may not intersect central axis <NUM>. For example, as illustrated in <FIG>, horizontal reference line <NUM> is positioned above central axis <NUM> along the vertical direction V. However, the circumference <NUM> of cross section <NUM> is once again constant. Such a configuration may be used to provide a more uniform flow distribution on boundary layer airflow circumferentially around fan inlet <NUM>.

Referring again to <FIG>, top half <NUM> of cross section <NUM> may have a maximum displacement <NUM> relative to reference circle <NUM>. According to the illustrated embodiment, the point of maximum displacement <NUM> of top half <NUM> of cross section <NUM> is at approximately <NUM> degrees and <NUM> degrees about the circumferential direction relative to the vertical direction V. Similarly, bottom half <NUM> of cross section <NUM> may have a maximum displacement <NUM> relative to reference circle <NUM>. According to the illustrated embodiment, the point of maximum displacement <NUM> of bottom half <NUM> of cross section <NUM> is at approximately <NUM> degrees and <NUM> degrees about the circumferential direction C<NUM> relative to the vertical direction V. It should be appreciated that these angles of maximum displacement are only approximates and may vary depending on the application. According to the illustrated embodiment, the maximum displacement <NUM> of top half <NUM> is equivalent to the maximum displacement <NUM> of bottom half <NUM>.

As illustrated in <FIG>, according to an exemplary embodiment, bottom half <NUM> of cross section may be tapered inward relative to reference circle <NUM>. More specifically, as illustrated, each side of bottom half <NUM> is tapered along a substantially straight line between the seven o'clock and the nine o'clock positions along the circumferential direction C, relative to a vertical reference line (not shown). However, according to alternative embodiments, bottom half <NUM> may take any shape suitable for improving the amount of boundary layer air to enter aft engine <NUM>.

An aircraft having a fuselage shaped in the manner described above and/or an aft engine configured in the manner described above may allow for capturing an optimal amount and distribution of a flow of boundary layer air from the fuselage. More specifically, the shaping of fuselage <NUM> results in a more uniform distribution of boundary layer airflow along the circumferential direction C<NUM> of the fuselage <NUM> and fan inlet <NUM>. The velocity of the boundary layer air flowing into the aft engine <NUM> may be similar from top half <NUM> to bottom half <NUM>, thus improving propulsive efficiency while reducing vibration, noise, and wear on the plurality of fan blades <NUM>.

Claim 1:
An aircraft (<NUM>) defining a longitudinal direction, a vertical direction, and a lateral direction, the aircraft (<NUM>) comprising:
a fuselage (<NUM>) extending between a forward end and an aft end along the longitudinal direction, the fuselage (<NUM>) defining a surface (<NUM>, <NUM>);
a stabilizer (<NUM>) attached to the fuselage (<NUM>) and extending between a leading edge (<NUM>) and a trailing edge (<NUM>);
a boundary layer ingestion fan (<NUM>) mounted to the fuselage (<NUM>), the boundary layer ingestion fan (<NUM>) defining a centerline (<NUM>) and comprising a plurality of fan blades (<NUM>) rotatable about the centerline (<NUM>) and a nacelle (<NUM>) surrounding the plurality of fan blades (<NUM>);
wherein the surface (<NUM>, <NUM>) of the fuselage (<NUM>) defines:
a first point (<NUM>) located in a plane perpendicular to the longitudinal direction and positioned where the leading edge (<NUM>) of the stabilizer (<NUM>) meets the fuselage (<NUM>);
a second point (<NUM>) located in a plane perpendicular to the longitudinal direction and positioned where the trailing edge (<NUM>) of the stabilizer (<NUM>) meets the fuselage (<NUM>);
an inflection point (<NUM>, <NUM>);
characterized by:
the boundary layer ingestion fan (<NUM>) being mounted at the aft end of the fuselage (<NUM>);
the fuselage further defining:
a first portion (<NUM>, <NUM>) of the surface, the first portion extending between the first point and the inflection point, the first portion being convex; and
a second portion (<NUM>, <NUM>) of the surface, the second portion extending between the inflection point and the second point, the second portion being concave.