Patent Description:
During operation of aircraft engines, an inlet of the aircraft engine can see a strong swirl distortion (or other airflow distortion) due to a variety of factors. For example in supersonic aircraft engines, the inventors of the present disclosure have found that relatively thin lips/ leading edges of a nacelle located upstream of a gas turbine engine can generate airflow distortion at the inlet to the gas turbine engine at certain aircraft operating conditions. The airflow distortion can be detrimental to an operability of the engine, and particularly to the gas turbine engine. Further, such airflow distortion can cause aeromechanical and/or operational issues.

Thus, an improved inlet assembly for an aircraft engine that addresses the aforementioned issue would be welcomed in the art. <CIT> discloses an inlet assembly for an aft fan for an aircraft. <CIT> discloses an inlet nacelle for a supersonic aircraft.

In one aspect, the invention is directed towards an engine according to claim <NUM>.

Generally, the present disclosure is directed to a low-distortion inlet assembly for reducing airflow swirl distortion entering an engine (or an aspect of the engine) mounted to or within an aircraft. Further, the inlet assembly includes a plurality of structural members (e.g. inlet guide vanes, struts, or similar) mounted at one or more predetermined locations around a circumference of an axis of the engine and at least one airflow modifying element configured within an inlet of the engine. More specifically, the predetermined locations have a distortion exceeding a predetermined threshold. As such, the inlet assembly is configured to reduce airflow distortion entering the engine or aspect of the engine.

The inlet guide vanes can be tailored to reduce flow distortion by introducing variations of the vanes. For example, in one embodiment, part-circumference inlet guide vanes may be located in groups at certain locations around an annulus where distortion is highest. In addition, one or more of the inlet guide vanes may be replaced with struts that provide structural support and flow turning to counter distortion. The inlet assembly of the present disclosure may also incorporate airflow modifying elements, such as vortex generators, trailing edge blowing, trailing edge suction, and/or high lift devices such as flaps attached to the structural members to further reduce distortion. Further, the inlet assembly of the present disclosure may also include non-axisymmetric internal area ruling or contouring to induce a flow field that counters the airflow distortion or moves it radially further away from the tips of the downstream compressor blades. Thus, the present invention reduces airflow distortion entering the engine or an aspect of the engine and reduces weight and helps improve the operability of the engine.

Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures, <FIG> illustrates a top view of one example of the aircraft <NUM> according to the present disclosure. <FIG> illustrates a port side view of the aircraft <NUM> as illustrated in <FIG>. As shown in <FIG> and <FIG> collectively, the aircraft <NUM> defines a longitudinal centerline <NUM> that extends therethrough, a vertical direction V, a lateral direction L, a forward end <NUM>, and an aft end <NUM>. As will be appreciated, the aircraft depicted in <FIG> and <FIG> may be a subsonic commercial aircraft, configured to operate at subsonic flight speeds.

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 <NUM> of the aircraft <NUM>. The first of such wings <NUM> extends laterally outwardly with respect to the longitudinal centerline <NUM> from a port side <NUM> of the fuselage <NUM> and the second of such wings <NUM> extends laterally outwardly with respect to the longitudinal centerline <NUM> from a starboard side <NUM> of the fuselage <NUM>. Further, as shown in the illustrated example, each of the wings <NUM> depicted includes one or more leading edge flaps <NUM> and one or more trailing edge flaps <NUM>. The aircraft <NUM> may also include 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. It should be appreciated however, that in other exemplary embodiments of the present disclosure, the aircraft <NUM> may additionally or alternatively include any other suitable configuration of stabilizer that may or may not extend directly along the vertical direction V or horizontal/ lateral direction L.

In addition, the aircraft <NUM> of <FIG> and <FIG> includes a propulsion system <NUM>, herein referred to as "system <NUM>. " The system <NUM> includes a pair of aircraft engines, at least one of which mounted to each of the pair of wings <NUM>, and an aft engine. For example, as shown, the aircraft engines are configured as turbofan jet engines <NUM>, <NUM> suspended beneath the wings <NUM> in an under-wing configuration. Additionally, the aft engine is configured as an engine that ingests and consumes air forming a boundary layer over the fuselage <NUM> of the aircraft <NUM>. Specifically, the aft engine is configured as a fan, i.e., a Boundary Layer Ingestion (BLI) fan <NUM>, configured to ingest and consume air forming a boundary layer over the fuselage <NUM> of the aircraft <NUM>. Further, as shown in <FIG>, the BLI fan <NUM> is mounted to the aircraft <NUM> at a location aft of the wings <NUM> and/or the jet engines <NUM>, <NUM>, such that a central axis <NUM> extends therethrough. As used herein, the "central axis" refers to a midpoint line extending along a length of the BLI fan <NUM>. Further, for the illustrated example, the BLI fan <NUM> is fixedly connected to the fuselage <NUM> at the aft end <NUM>, such that the BLI fan <NUM> is incorporated into or blended with a tail section at the aft end <NUM>. However, it should be appreciated that in various other embodiments, some of which will be discussed below, the BLI fan <NUM> may alternatively be positioned at any suitable location of the aft end <NUM>.

In various embodiments, the jet engines <NUM>, <NUM> may be configured to provide power to an electric generator <NUM> and/or an energy storage device <NUM>. For example, one or both of the jet engines <NUM>, <NUM> may be configured to provide mechanical power from a rotating shaft (such as an LP shaft or HP shaft) to the electric generator <NUM>. Additionally, the electric generator <NUM> may be configured to convert the mechanical power to electrical power and provide such electrical power to one or more energy storage devices <NUM> and/or the BLI fan <NUM>. Accordingly, in such examples, the propulsion system <NUM> may be referred to as a gas-electric propulsion system. It should be appreciated, however, that the aircraft <NUM> and propulsion system <NUM> depicted in <FIG> and <FIG> is provided by way of example only and that in other examples of the present disclosure, any other suitable aircraft <NUM> may be provided having a propulsion system <NUM> configured in any other suitable manner.

Referring now to <FIG>, in certain examples, the jet engines <NUM>, <NUM> may be configured as high-bypass turbofan jet engines. More specifically, <FIG> illustrates a schematic cross-sectional view of one example of a high-bypass turbofan jet engine <NUM>, herein referred to as "turbofan <NUM>. " In various examples, the turbofan <NUM> may be representative of jet engines <NUM>, <NUM>. Further, as shown, the turbofan <NUM> 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>.

In particular examples, the core turbine engine <NUM> generally includes a substantially tubular outer casing <NUM> that defines an annular inlet <NUM>. It should be appreciated, that as used herein, terms of approximation, such as "approximately," "generally," "substantially," or "about," refer to being within a forty percent margin of error. 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>.

Further, as shown, 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 the disk <NUM> generally along the radial direction R<NUM>. Each fan blade <NUM> is rotatable relative to the disk <NUM> about a pitch axis 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. As such, the fan blades <NUM>, the disk <NUM>, and the actuation member <NUM> are together rotatable about the longitudinal axis <NUM> by LP shaft <NUM> across a power gearbox <NUM>. In certain examples, the power gearbox <NUM> includes a plurality of gears for stepping down the rotational speed of the LP shaft <NUM> to a more efficient rotational fan speed.

Referring still to <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 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 outer 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.

In addition, it should be appreciated that the turbofan engine <NUM> depicted in <FIG> is by way of example only, and that in other examples, the turbofan engine <NUM> may have any other suitable configuration. Further, it should be appreciated, that in other examples, the jet engines <NUM>, <NUM> may instead be configured as any other suitable aeronautical engine.

Referring now to <FIG>, a schematic, cross-sectional side view of an aft engine in accordance with various examples of the present disclosure is provided, such as the aft engine mounted to the aircraft <NUM> at the tail section <NUM> of the aircraft <NUM>. More specifically, as shown, the aft engine is configured as a boundary layer ingestion (BLI) fan <NUM>. The BLI fan <NUM> may be configured in substantially the same manner as the BLI fan <NUM> described above with reference to <FIG> and <FIG> and the aircraft <NUM> may be configured in substantially the same manner as the exemplary aircraft <NUM> described above with reference to <FIG> and <FIG>.

More specifically, as shown, the BLI fan <NUM> defines an axial direction A<NUM> extending along the central axis <NUM> that extends therethrough for reference. Additionally, the BLI fan <NUM> defines a radial direction R<NUM> and a circumferential direction (not shown). In general, the BLI fan <NUM> includes a fan <NUM> rotatable about the central axis <NUM>, a nacelle <NUM> extending around at least 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>. Further, the fan <NUM> includes a plurality of fan blades <NUM> spaced generally along the circumferential direction C<NUM>. Moreover, the structural member(s) <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>. Additionally, the nacelle <NUM> extends around and encircles the plurality of fan blades <NUM>, and also extends around the fuselage <NUM> of the aircraft <NUM> at an aft end <NUM> of the aircraft <NUM> when, as shown in <FIG>, the BLI fan <NUM> is mounted to the aircraft <NUM>. Notably, as used herein, the term "nacelle" includes the nacelle as well as any structural fan casing.

As is also depicted in <FIG>, the fan <NUM> additionally 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 examples, 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). Changing the pitch of the plurality of fan blades <NUM> may increase an efficiency of the BLI fan <NUM> and/or may allow the BLI fan <NUM> to achieve a desired thrust profile. With such an example, the BLI fan <NUM> may be referred to as a variable pitch BLI fan.

The fan shaft <NUM> is mechanically coupled to a power source <NUM> located at least partially within the fuselage <NUM> of the aircraft <NUM>, forward of the plurality of fan blades <NUM>. Further, as shown, 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 BLI fan <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. With such an example, the gearbox <NUM> may be operably connected to, e.g., a controller of the aircraft <NUM> for changing its ratio in response to one or more flight conditions.

In certain examples, the BLI fan <NUM> may be configured with a gas-electric propulsion system, such as the gas-electric propulsion system <NUM> described above with reference to <FIG>. In such an example, the power source <NUM> may be an electric motor that receives power from one or both of an energy storage device or an electric generator- such as the energy storage device <NUM> or electric generator <NUM> of <FIG> and <FIG>, the electric generator <NUM> converting mechanical power received from one or more under-wing mounted aircraft engines to electric power. However, in other examples, the power source <NUM> may instead be any other suitable power source. For example, the power source <NUM> may alternatively be configured as a gas engine, such as a gas turbine engine or internal combustion engine. Moreover, in certain examples, the power source <NUM> may be positioned at any other suitable location within, e.g., the fuselage <NUM> of the aircraft <NUM> or the BLI fan <NUM>. For example, in certain examples, the power source <NUM> may be configured as a gas turbine engine positioned at least partially within the BLI fan <NUM>.

As briefly stated above, the BLI fan <NUM> includes one or more structural members <NUM> for mounting the BLI fan <NUM> to the aircraft <NUM>. More specifically, as shown in <FIG>, the structural member(s) <NUM> may be configured as inlet guide vanes <NUM> for the fan <NUM> and/or as struts <NUM>. Further, it should be understood that the structural member(s) <NUM> may be configured as fixed inlet guide vanes extending between the nacelle <NUM> and the fuselage <NUM> of the aircraft <NUM>. Alternatively, the structural member(s) <NUM> may be configured as variable inlet guide vanes. Further, as shown, the structural member(s) <NUM> generally extend substantially along the radial direction R<NUM> of the BLI fan <NUM> between the nacelle <NUM> and the fuselage <NUM> of the aircraft <NUM> for mounting the BLI fan <NUM> to the fuselage <NUM> of the aircraft <NUM>. As such, the structural member(s) <NUM> may be shaped and/or oriented to direct and/or condition a flow of air into the BLI fan <NUM> to, e.g., increase an efficiency of the BLI fan <NUM>, or reduce a distortion of the air flowing into the BLI fan <NUM>, which will be discussed in more detail below.

Referring still to <FIG>, the BLI fan <NUM> additionally includes one or more outlet guide vanes <NUM> and a tail cone <NUM>. The one or more outlet guide vanes <NUM> for the example depicted extend between the nacelle <NUM> and the tail cone <NUM> for directing a flow of air through the BLI fan <NUM>, and optionally for adding strength and rigidity to the BLI fan <NUM>. The outlet guide vanes <NUM> may be evenly spaced along the circumferential direction C<NUM> 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. Inclusion of the plurality of outlet guide vanes <NUM> extending between the nacelle <NUM> and the tail cone <NUM> allows for maximizing the efficiency of the BLI fan <NUM>.

Further, aft of the plurality of fan blades <NUM>, and for the example depicted, aft of the one or more outlet guide vanes <NUM>, the BLI fan <NUM> additionally defines a nozzle <NUM> between the nacelle <NUM> and the tail cone <NUM>. As such, the nozzle <NUM> may be configured to generate an amount of thrust from the air flowing therethrough. In addition, the tail cone <NUM> may be shaped to minimize an amount of drag on the BLI fan <NUM>. However, in other examples, 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 examples, the BLI fan <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 particularly to <FIG> and <FIG>, the BLI fan <NUM> defines an inlet <NUM> at a forward end <NUM> between the nacelle <NUM> and the fuselage <NUM> of the aircraft <NUM>. As mentioned above, the nacelle <NUM> of the BLI fan <NUM> extends around the central axis <NUM> of the aircraft <NUM> and the fuselage <NUM> of the aircraft <NUM> at the aft end of the aircraft <NUM>. Accordingly, as shown, the inlet <NUM> of the BLI fan <NUM> extends substantially three hundred sixty degrees (<NUM>°) around the central axis <NUM> of the aircraft <NUM> and the fuselage <NUM> of the aircraft <NUM> when, such as in the example depicted, the BLI fan <NUM> is mounted to the aircraft <NUM>. Additionally, in still further examples, the BLI fan <NUM>, or rather the external surface of the nacelle <NUM>, may have any other suitable cross-sectional shape along the axial direction A<NUM> (as opposed to the circular shape depicted) and the structural members <NUM> may not be evenly spaced along the circumferential direction C<NUM>.

Referring particularly to <FIG>, a schematic, cross-sectional view of one example of the BLI fan <NUM>, viewed along an axial centerline <NUM> thereof so as to illustrate an inlet assembly <NUM> according to the present disclosure is illustrated. More specifically, as shown, the illustrated BLI fan <NUM> includes a plurality of structural members <NUM> mounted at one or more predetermined locations around a circumference of the fan shaft <NUM> of the BLI fan <NUM>. For example, in certain examples, the predetermined locations as described herein have a distortion exceeding a predetermined threshold. In other words, for certain examples, the airflow entering the BLI fan <NUM> may be evaluated to determine a pattern thereof. Thus, the location and/or number of structural members <NUM>, as well as the shape of the structural members <NUM>, may be designed and chosen as a function of the pattern or distortion. In further examples, as shown in <FIG>, only a portion of the circumference of the fan shaft <NUM> may include structural members <NUM>. For example, in the example shown, about fifty percent (<NUM>%) of the circumference includes structural member <NUM>. Alternatively, the structural members <NUM> may be spaced around the entire circumference of the fan shaft <NUM>. As such, in particular examples, the structural members <NUM> may be evenly spaced along the circumferential direction C<NUM> of the BLI fan <NUM> around the fan shaft <NUM>. In alternative examples, the structural members may form one or more inlet guide vane groups <NUM>, which may be spaced appropriately depending on the distortion pattern along the circumferential direction C<NUM> of the BLI fan <NUM> around the fan shaft <NUM>.

In addition, as shown in the illustrated example, the structural members <NUM> may be located circumferentially at a substantially twelve o'clock, a substantially three o'clock, a substantially six o'clock, and/or a substantially nine o'clock, receptively, with respect to the circumference of the fan shaft <NUM>. It should be understood that the predetermined locations may be at the illustrated locations as well as any location therebetween and are meant to encompass locations having a high distortion and/or a location where a modification of the airflow would have the highest impact of correcting the distortion. Further, as mentioned, the structural members <NUM> may include inlet guide vanes <NUM>, struts <NUM>, or similar or any combinations thereof.

Still referring to <FIG>, the inlet assembly <NUM> may also include at least one airflow modifying element <NUM> configured within the inlet <NUM>. As such, the inlet assembly <NUM> of the present disclosure is configured to reduce distortion of the airflow entering the fan <NUM>. Accordingly, the present disclosure may include any suitable combination of structural members <NUM> and/or airflow modifying elements <NUM> so as to offset or modify the distortion of the airflow entering the BLI fan <NUM>, examples of which are described in more detail below.

For example, as shown in <FIG>, the inlet assembly <NUM> may include a plurality of inlet guide vanes <NUM> placed in groups <NUM> at the predetermined locations around the circumference of the fan shaft <NUM> (i.e. spaced along the circumferential direction C<NUM> of the BLI fan <NUM>), each extending between the nacelle <NUM> and the fuselage <NUM> of the aircraft <NUM> generally along the radial direction R<NUM>. For example, as shown, the inlet assembly <NUM> includes a single group <NUM> of three inlet guide vanes <NUM>. In alternative examples, the inlet assembly <NUM> may include more than one group <NUM> of inlet guide vanes <NUM> at any circumferential location having any suitable number of inlet guide vanes <NUM>. In addition, for examples having more than one inlet guide vane group <NUM>, each group <NUM> may include the same number of inlet guide vanes <NUM> or a different number of inlet guide vanes <NUM>. In further examples, the inlet assembly <NUM> may include a plurality of inlet guide vanes <NUM> grouped into a plurality of separate and distinct inlet guide vane groups <NUM> around the circumference of the fan shaft <NUM>.

In addition, the inlet assembly <NUM> may include one or more struts <NUM> extending between the nacelle <NUM> and the fuselage <NUM> of the aircraft <NUM> generally along the radial direction R<NUM>. Generally, struts are structural components designed to resist longitudinal compression. In addition, the struts <NUM> of the present disclosure are strategically placed at the predetermined locations so as to redistribute the airflow entering the fan <NUM> more uniformly circumferentially so as to reduce swirl distortion at the inlet <NUM>. For example, as shown, the illustrated inlet assembly <NUM> includes at least two struts <NUM>, i.e. at the three o'clock and the six o'clock positions, respectively. In further examples, the inlet assembly <NUM> may include more than two or less than two struts <NUM>.

Referring particularly to <FIG>, the airflow modifying element(s) <NUM> may include at least one of a vortex generator <NUM> (<FIG>), a contoured surface <NUM> (<FIG>), a flap <NUM> (<FIG>), or any other airflow modifying element. More specifically, as shown in <FIG>, the inlet assembly <NUM> may include a contoured surface <NUM> or area ruling on an inner surface <NUM> of the nacelle <NUM>, e.g. at the twelve o'clock position, that is configured to push airflow radially inward and away from the contoured surface <NUM>. As such, the contouring of the inner surface <NUM> of the nacelle <NUM> is configured to induce a flow field that counters a type of swirl distortion of the airflow entering the inlet <NUM>. In addition, the inner surface <NUM> of the nacelle <NUM> may also include one or more indentions configured to modify the airflow.

Further, as shown in <FIG>, one or more of the structural members <NUM> may include a vortex generator <NUM> and/or a flap <NUM> configured therewith. More specifically, <FIG> illustrate cross-sectional views of an inlet guide vane <NUM> taken along the radial direction R<NUM> that may be included in the inlet assembly <NUM>. As shown, the inlet guide vane <NUM> extends between a forward, upstream end <NUM> and an aft, downstream end <NUM>. The forward, upstream end <NUM> includes a leading edge <NUM> of the inlet guide vane <NUM> and the aft, downstream end <NUM> includes a trailing edge <NUM> of the inlet guide vane <NUM>. A body <NUM> of the inlet guide vane <NUM> is fixed relative to the nacelle <NUM> of the BLI fan <NUM> and the fuselage <NUM> of the aircraft <NUM> and includes a pressure side <NUM> and a suction side <NUM>. Further, as shown, the inlet guide vane <NUM> may include one or more vortex generators <NUM> configured on one or more of the pressure or suction sides <NUM>, <NUM>. For example, as shown, the illustrated inlet guide vane <NUM> includes a single vortex generator <NUM> on the suction side <NUM> thereof. In additional examples, the inlet guide vanes <NUM> and/or the struts <NUM> may include any number and/or type of vortex generators or similar surface features mounted to a surface thereof so as to redirect the airflow entering the BLI fan <NUM>.

In addition, as shown in <FIG>, the inlet guide vanes <NUM> may also include an optional flap <NUM> at the aft end <NUM> configured to rotate about a substantially radial axis <NUM>. For example, as shown, the flap <NUM> is configured to rotate between a first position <NUM> (in phantom), a neutral position <NUM>, a second position <NUM> (in phantom), and a potentially infinite number of positions therebetween. By rotating the flap <NUM> between the various positions, the inlet guide vanes <NUM> may be configured to vary a direction in which air flowing thereover is directed.

In yet another example, the inlet assembly <NUM> may include trailing edge blowing or suction that is configured to reduce axial or swirl distortion entering the BLI fan <NUM>. In addition, the inlet assembly <NUM> may include angled flow injection. Generally, trailing edge blowing encompasses flow injection along the direction of the airflow. In contrast, angled flow injection encompasses flow injection at an angle. Further, the flow injection may be steady or unsteady. As used herein, trailing edge blowing generally refers to a technique of injecting air into the inlet <NUM> at or near the trailing edge <NUM> of the inlet guide vanes <NUM> or slightly upstream of the trailing edge <NUM>. For example, in one example, trailing edge blowing may include injecting airflow into the main airstream through a hole or slot configured within the airfoil. As used herein, trailing edge suction generally refers to a technique draining air from the inlet <NUM> at or near the trailing edge <NUM> of the inlet guide vanes <NUM> or slightly upstream of the trailing edge <NUM>. As such, both trailing edge blowing or trailing edge suction are configured to modify the airflow entering the inlet <NUM> so as to reduce airflow distortion entering the fan <NUM>. Further, trailing edge blowing can be achieved by steady or pulsed blowing aligned with the airflow or at an angle to achieve the same effect as a miniature vortex generator or tab. Notably, as used herein, the term "airflow distortion" refers to variation in airflow properties, such airflow properties including airflow speed, airflow pressure, etc. Accordingly, airflow distortion entering the fan <NUM> refers to variations in these airflow properties over an entire face of the fan (circumferentially and radially), at a location downstream of the inlet and upstream of the fan, such as at a location immediately upstream of the fan.

Referring now to <FIG>, cross-sectional views of additional examples of the inlet guide vanes <NUM> of the present disclosure are illustrated. It should be understood that such features can also be applied to struts. As shown, each of the inlet guide vanes <NUM> may have a unique shape and/or orientation corresponding to airflow conditions entering the BLI fan <NUM> at a particular location in the fan <NUM>. Thus, any combination of shapes may be used in the inlet assembly <NUM> and can be chosen based on a determined swirl distortion of the airflow entering the BLI fan <NUM>.

For example, as shown, each of the inlet guide vanes <NUM> may have a cambered upright airfoil cross-section (<FIG>), a cambered inverted airfoil cross-section (<FIG>), or a symmetrical airfoil cross-section (<FIG>). More specifically, as shown in <FIG>, the cambered upright inlet guide vane <NUM> generally has a mean camber line <NUM> above the chord line <NUM> of the airfoil, with the trailing edge <NUM> having a downward direction. Such cambered airfoils typically generate lift at zero angle of attack and since air follows the trailing edge <NUM>, the air is deflected downward. As shown in <FIG>, the inverted inlet guide <NUM> vane generally has a mean camber line <NUM> below the chord line <NUM> of the airfoil, with the trailing edge <NUM> having an upward direction. When a cambered airfoil is upside down, the angle of attack can be adjusted so that the lift force is upwards. In contrast, as shown in <FIG>, the mean camber line <NUM> and the chord line <NUM> of a symmetrical airfoil are the same (i.e. the lines <NUM>, <NUM> overlap and there is zero chamber).

It should be understood that the lift force depends on the shape of the airfoil, especially the amount of camber (i.e. curvature such that one surface is more convex than the other surface). In other words, increasing the camber of the airfoil turns the flow more which in turn generally increases lift. The local turning of the flow can be used to counter the local flow distortion and result in a more uniform flow profile ingested by the fan. Additionally, or alternatively, the flow turning may be used to generate a favorable swirl profile entering the fan radially, helping improve an efficiency and operability of the fan under distortion.

In addition, as shown generally in <FIG>, a leading edge radius <NUM> of one or more of the inlet guide vanes <NUM> may be designed to reduce swirl distortion of the airflow entering the BLI fan <NUM>. In addition, as shown in <FIG>, in certain examples, the leading edge radius <NUM> of one or more of the inlet guide vanes <NUM> may vary in a span-wise direction <NUM> (e.g. get larger or smaller) as a function of the airflow conditions entering the BLI fan <NUM>. As such, the leading edge radius <NUM> of each inlet guide vane <NUM> can be designed according to the flow conditions it receives. In addition, as shown, a camber angle (i.e. a curve) for an individual vane <NUM> can also vary in the span-wise direction <NUM> to effectively turn the flow to a more uniform state at the discharge of the inlet guide vane <NUM>.

It will be appreciated, however, that the vanes <NUM> depicted in <FIG> are by way of example only, and are depicted schematically for exemplary purposes. In certain examples, the vanes <NUM> may each be thinner than depicted, particularly when incorporated into an engine of supersonic aircraft. Further, it will be appreciated that in other examples, the inlet assembly <NUM> may be incorporated into any other suitable engine for any other suitable aircraft. For example, in certain examples, the inlet assembly <NUM> may be incorporated into, or otherwise operable with, any other suitable gas turbine engine, such as a supersonic gas turbine engine for mounting in or to a supersonic aircraft. In certain examples, the engine may be mounted at an aft end of the aircraft (similar to engine <NUM> of aircraft <NUM>), may be mounted on or to the wings, may be incorporated into the body of the aircraft (e.g., incorporated into the wings, fuselage, stabilizer, etc.), or may be operable with the aircraft in any other suitable manner.

More specifically, referring now to <FIG>, an engine <NUM> in accordance with another exemplary aspect of the present disclosure is provided. As shown, the engine <NUM> generally includes an engine casing <NUM> and a gas turbine engine <NUM>. For the embodiment shown, the engine casing <NUM> is configured as a nacelle surrounding at least in part the gas turbine engine <NUM>. The gas turbine engine <NUM> generally includes a stage of compression airfoils and a turbine coupled to the stage of compression airfoils through an engine shaft for driving the stage of compression airfoils. More particularly, for the embodiment shown, the gas turbine engine <NUM> includes a compressor section having a compressor and a turbine section having the turbine, and the stage of compression airfoils is a stage of compressor rotor blades of the compressor of the compressor section of the gas turbine engine <NUM>. However, in other embodiments, the stage of compression airfoils may instead be, e.g., a plurality of fan blades of a fan (see, e.g., fan blades <NUM> of <FIG>).

Referring still to the gas turbine engine <NUM> of <FIG>, the compressor is a first compressor <NUM>, and the exemplary gas turbine engine <NUM> depicted further includes a second compressor <NUM>. The first compressor <NUM> may be a low pressure compressor and the second compressor <NUM> may be a high pressure compressor. Moreover, the turbine section includes a first turbine <NUM> and a second turbine <NUM>. The first turbine <NUM> may be a high pressure turbine coupled to the high pressure compressor/ second compressor <NUM>, and the second turbine <NUM> may be a low pressure turbine coupled to the low pressure compressor/ first compressor <NUM>. A combustion section <NUM> is located between the second compressor <NUM> and the first turbine <NUM>. Further, the gas turbine engine <NUM> defines a nozzle <NUM> downstream of the turbine section.

As is also depicted, the gas turbine engine <NUM> includes a nose cone <NUM> and a gas turbine engine casing <NUM>, with the gas turbine engine casing <NUM> surrounding the compressor section and the turbine section and defining an inlet <NUM>. As also noted above, the engine <NUM> includes the engine casing <NUM>, also referred to as the nacelle. The nacelle defines an airflow duct <NUM> upstream of the gas turbine engine inlet <NUM>, and further defines a nacelle inlet <NUM> upstream of the airflow duct <NUM>. More specifically, the casing <NUM> includes a forward lip <NUM> defining the nacelle inlet <NUM>. As indicated in the close-up Callout Circle A of <FIG>, the forward lip <NUM> defines a relatively small radius of curvature <NUM>, such as a radius of curvature less than about two inches, such as less than about <NUM> inch, such as less than about <NUM> inches, such as less than about <NUM> inches. For the embodiment shown, the nacelle inlet <NUM> defines an angle <NUM> with the radial direction R (depicted relative to reference line R' in <FIG>) greater than about fifteen degrees, such as greater than about twenty degrees, such as greater than about twenty-five degrees, such as greater than about thirty degrees, and up to about eighty-five degrees. Moreover, for the embodiment shown, the airflow duct <NUM>, which extends between the gas turbine engine inlet <NUM> and the nacelle inlet <NUM>, defines a centerline <NUM> (i.e., a reference line extending halfway between a top wall and a bottom wall in the plane depicted). For the embodiment shown, the centerline <NUM> is a non-linear centerline <NUM>. It will further be appreciated that the nacelle defines a bypass passage <NUM> around the gas turbine engine <NUM>, which may be useful during certain phases of flight of a supersonic aircraft. The nonlinear centerline <NUM> of the airflow duct <NUM>, may assist with facilitating an airflow into and through the bypass passage <NUM>.

However, such features, such as the sharp nacelle lip <NUM>, may create an airflow distortion when the aircraft is subject to an angle of attack at an upstream-most stage of compression airfoils of the gas turbine engine <NUM>, or more particularly, at a first stage of compressor rotor blades <NUM> of the first compressor <NUM> of the compressor section of the gas turbine engine <NUM>. Accordingly, the exemplary engine <NUM> also includes a low distortion inlet assembly <NUM> mounted within the inlet <NUM>. Referring now briefly also to <FIG>, providing a close-up view of the exemplary low distortion inlet assembly <NUM>, viewed along a central axis <NUM> of the engine <NUM>, it will be appreciated that the exemplary low distortion inlet assembly <NUM> may be configured in a similar manner to the exemplary low distortion inlet assembly <NUM> described above with reference to <FIG>. The same numbers accordingly refer to the same parts.

As shown, it will be appreciated that the low distortion inlet assembly <NUM> may generally include one or more structural members <NUM> mounted at predetermined locations around the circumference of a central axis <NUM> of the engine <NUM> within the inlet <NUM> defined by the gas turbine engine casing <NUM> (e.g., extending between the gas turbine engine casing <NUM> and the nose cone <NUM>). The predetermined locations define an airflow distortion exceeding a predetermined threshold. The low distortion inlet assembly <NUM> may also include at least one airflow modifying element <NUM> configured within the inlet <NUM> so as to reduce airflow distortion into the stage of compression airfoils (which, as noted above, is configured as a stage of compressor rotor blades <NUM>).

As will be appreciated, the low distortion inlet assembly <NUM> may assist with accommodating, or correcting, a distortion in the airflow into the gas turbine engine <NUM> resulting from certain features of the supersonic engine <NUM>. The inlet assembly <NUM> of <FIG> and <FIG> may include one or more of the features described above with respect to the inlet assembly <NUM> incorporated into the aft-mounted engine. For example, the inlet assembly <NUM> of <FIG> and <FIG> may include one or more of the structural members <NUM> and airflow modifying elements <NUM> described above with reference to <FIG> (except incorporated into an inlet <NUM> of a gas turbine engine <NUM> of a supersonic engine <NUM>).

By further way of example, referring to <FIG>, a close-up view is provided of a low distortion inlet assembly <NUM> in accordance with an embodiment of the present disclosure, viewed along a central axis <NUM> of an engine <NUM>. The inlet assembly <NUM> depicted may be configured in a similar manner as the exemplary inlet assembly <NUM> of <FIG>. For example, the inlet assembly <NUM> includes a plurality of structural members <NUM> mounted at one or more predetermined locations around a circumferential direction C of the inlet assembly <NUM>. As with the embodiments above, the placement of the structural members <NUM> may be set to address an airflow distortion through the inlet assembly <NUM> (e.g., through an inlet <NUM> of the engine casing <NUM>/ nacelle if the inlet assembly <NUM> is mounted within the airflow duct <NUM> of the casing <NUM>, and/or through the inlet <NUM> of the gas turbine engine <NUM> if the inlet assembly <NUM> is mounted at the inlet <NUM> of the gas turbine engine <NUM>).

More specifically, for the embodiment shown, the structural members <NUM> are asymmetrically spaced along the circumferential direction C, with a density of the structural members <NUM> being higher where a higher airflow distortion is expected. For example, in certain embodiments, the structural members may form one or more structural member groups <NUM>, which may be spaced appropriately depending on a distortion pattern along the circumferential direction C. Additionally, or alternatively, a shape of the structural members <NUM> may be designed and chosen as a function of the pattern of distortion (see, e.g., <FIG>). For the embodiment shown the inlet assembly <NUM> includes two structural member groups <NUM> positioned substantially in a top half. More specifically, for the embodiment shown, the inlet assembly <NUM> defines a first circumferential portion having a first density of structural members <NUM> and a second circumferential portion having a second density of structural members <NUM>, and the first density is different than the second density. However, in other embodiments, the inlet assembly <NUM> may include any other suitable number of groups <NUM>, positioned at any suitable location circumferentially.

For the embodiment of <FIG>, the structural members <NUM> are inlet guide vanes <NUM>, but they may additionally or alternatively be configured as struts <NUM>. It should be understood that the predetermined locations may be at the illustrated locations as well as any location therebetween and are meant to encompass locations having a high distortion and/or a location where a modification of the airflow would have the highest impact of correcting the distortion. Moreover, in certain embodiments, the inlet assembly <NUM> may be mounted at the inlet <NUM> to the gas turbine engine <NUM> (i.e., within the casing <NUM>), or alternatively may be mounted within the engine casing <NUM> of the engine <NUM>, e.g., within the inlet duct <NUM>.

Further, it should be appreciated that the inlet assembly <NUM> and engine <NUM> discussed above are by way of example only. In other embodiments the airflow duct <NUM> of the casing <NUM> may have any other suitable shape having a non-linear centerline (e.g., a serpentine airflow duct <NUM>). More specifically, <FIG> shows a forward end of an engine <NUM> having a casing <NUM> in accordance with another embodiment of the present disclosure. The engine <NUM> further includes a nosecone <NUM> having a more streamlined shape. Further the nosecone <NUM> depicted may extend forward of an inlet <NUM> defined by the casing <NUM> (as in the embodiment shown), but alternatively may not extend forward of the inlet <NUM>. Further, for the embodiment shown in <FIG>, the inlet assembly <NUM> is positioned within the airflow duct <NUM> of the casing <NUM>, at a location upstream of the inlet <NUM> to the gas turbine engine <NUM> (not shown in <FIG>) of the engine <NUM>.

Claim 1:
A turbofan or turbojet gas turbine engine (<NUM>) for mounting in or to a supersonic aircraft, the gas turbine engine (<NUM>) defining a central axis (<NUM>), wherein the gas turbine engine (<NUM>) comprises:
a stage of compression airfoils rotatable about the central axis (<NUM>);
a turbine (<NUM>, <NUM>) coupled to the stage of compression airfoils for driving the stage of compression airfoils;
a casing (<NUM>) surrounding the stage of compression airfoils and defining an inlet (<NUM>), wherein the inlet (<NUM>) is positioned upstream of the stage of compression airfoils and wherein the casing (<NUM>) is a gas turbine engine casing;
an engine casing (<NUM>) surrounding the gas turbine engine , the engine casing configured as a nacelle; and
a low-distortion inlet assembly (<NUM>) mounted within the inlet (<NUM>), the low-distortion inlet assembly (<NUM>) comprising
one or more structural members (<NUM>) mounted at predetermined locations around a circumference of the central axis (<NUM>) within the inlet (<NUM>), the predetermined locations defining an airflow distortion exceeding a predetermined threshold; and
at least one airflow modifying element (<NUM>) configured within the inlet (<NUM>) so as to reduce airflow distortion entering the stage of compression airfoils;
characterized in that an airflow duct (<NUM>) of the nacelle defines a non-linear centerline (<NUM>).