Aircraft with rear engine and air injection assembly for the aircraft

Abstract

An air injection assembly for an aircraft is provided. The aircraft includes a fuselage extending longitudinally between a front end and a rear end, and a boundary layer intake fan mounted to the fuselage at the rear end of the fuselage. The air injection assembly includes a plurality of injection ports defined on a surface of the fuselage at locations upstream of the boundary layer intake fan. Supplemental airflow is provided to the injection ports through a fluid passageway, and the supplemental airflow is injected into the injection ports to displace at least a portion of the relatively high-velocity boundary layer airflow. In this manner, the airflow entering the boundary layer intake fan is more uniform, has less vortex distortion, and has a lower average velocity.

Description

[0001] This application is a divisional application of the invention patent application with application number 201780072546.X filed on August 25, 2017 and invention name “Aircraft with rear engine and air injection assembly for the aircraft”. Technical Field

[0002] The present subject matter generally relates to an aircraft having a rear engine, or more particularly to an aircraft fuselage having a rear engine for increasing the efficiency of the rear engine. Background Art

[0003] A conventional commercial aircraft typically 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 aircraft's wings, for example, in a suspended position below the wing, separate from the wing and fuselage. This configuration allows the turbofan jet engines to interact with a separate freestream airflow, unaffected by the wings and / or fuselage. This configuration can reduce the amount of turbulence within the air entering the inlet of each respective turbofan jet engine, which has a positive impact on the net thrust of the aircraft.

[0004] However, drag on an aircraft, including a turbofan jet engine, also affects the net thrust of the aircraft. The total drag on an aircraft, including surface friction and form drag, is generally proportional to the difference between the freestream velocity of the air approaching the aircraft and the average velocity of the wake downstream of the aircraft due to the drag on the aircraft.

[0005] Positioning a fan at the aft end of an aircraft fuselage can help re-energize the boundary layer airflow over the aft end of the fuselage. 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 consistent velocity profile along the circumference of the fan. More specifically, the structures at the aft end of the fuselage may generate wake vortices, resulting in vortex distortion and an inconsistent velocity profile along the circumference of the airflow ingested by the fan.

[0006] Therefore, an aircraft capable of encouraging slow-moving air to form a boundary layer across the aircraft fuselage would be useful. Specifically, an aircraft fuselage designed to increase the intake of relatively low-momentum boundary layer airflow entering the rear engine and reduce the non-uniformity and distortion of the velocity distribution of the intake airflow would be particularly beneficial. Summary of the Invention

[0007] Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention.

[0008] In one exemplary embodiment of the present disclosure, an aircraft is provided that defines a longitudinal direction, a vertical direction, and a lateral direction. The aircraft includes a fuselage extending along the longitudinal direction between a front end and a rear end, and a boundary layer intake fan mounted to the fuselage at the rear end of the fuselage, the boundary layer intake fan defining a centerline and including a plurality of fan blades capable of rotating about the centerline. An air injection assembly includes: a plurality of injection ports defined on a surface of the fuselage at locations upstream of the boundary layer intake fan; and a fluid passage extending through the fuselage and in fluid communication with the plurality of injection ports for directing a flow of supplemental air through the plurality of injection ports.

[0009] In another exemplary embodiment of the present disclosure, an air injection assembly for an aircraft is provided. The aircraft includes a fuselage extending in a longitudinal direction between a front end and a rear end, and a boundary layer intake fan mounted to the fuselage at the rear end of the fuselage. The air injection assembly includes: a plurality of injection ports defined on a surface of the fuselage at locations upstream of the boundary layer intake fan; and a fluid channel extending through the fuselage and in fluid communication with the plurality of injection ports for directing a supplemental airflow through the plurality of injection ports.

[0010] These and other features, aspects and advantages of the present invention will be better understood with reference to the following description and appended claims.The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the present invention and, together with the description, serve to explain the principles of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0011] This specification, with reference to the accompanying drawings, sets forth a complete and enabling disclosure of the invention, including the best mode thereof, to one of ordinary skill in the art.

[0012] Figure 1 is a top view of an aircraft according to various exemplary embodiments of the present disclosure.

[0013] Figure 2 yes Figure 1 A port side view of an exemplary aircraft.

[0014] Figure 3 is a schematic cross-sectional view of a gas turbine engine according to an exemplary embodiment of the present disclosure.

[0015] Figure 4 According to an exemplary embodiment of the present disclosure Figure 1 A schematic cross-sectional view of the aft end of the fuselage of an exemplary aircraft.

[0016] Figure 5 According to another exemplary embodiment of the present disclosure Figure 1 A schematic cross-sectional side view of the aft end of the fuselage of an exemplary aircraft.

[0017] Figure 6 According to another exemplary embodiment of the present disclosure Figure 1 A schematic cross-sectional side view of the aft end of the fuselage of an exemplary aircraft.

[0018] Figure 7 According to another exemplary embodiment of the present disclosure Figure 1 A schematic cross-sectional side view of the aft end of the fuselage of an exemplary aircraft. DETAILED DESCRIPTION

[0019] Reference will now be made in detail to embodiments of the invention, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. The same or similar designations in the drawings and the description have been used to refer to the same or similar parts of the invention. As used herein, the terms "first," "second," and "third" are used interchangeably to distinguish one component from another and are not intended to indicate the position or importance of the various components. The terms "upstream" and "downstream" refer to relative directions with respect to the flow of a fluid in a fluid path. For example, "upstream" refers to the direction from which a fluid is flowing and "downstream" refers to the direction toward which a fluid is flowing.

[0020] Referring now to the drawings, wherein like numerals refer to like elements throughout, Figure 1 A top view of an exemplary aircraft 10 is provided, which may incorporate various embodiments of the present invention. Figure 2 Provides such Figure 1 The port side 24 of the aircraft 10 is shown. Figure 1 and Figure 2 As shown in FIG, the aircraft 10 defines a longitudinal direction 12 extending therethrough, a vertical direction V, a lateral direction L, a front end 14, and a rear end 16. Furthermore, the aircraft 10 defines a mean line 18 extending between the front end 14 and the rear end 16 of the aircraft 10. As used herein, "mean line" refers to a midpoint line extending along the length of the aircraft 10 without taking into account appendages of the aircraft 10 (e.g., wings 22 and stabilizers discussed below).

[0021] Furthermore, aircraft 10 includes a fuselage 20 and a pair of wings 22 extending longitudinally from a front end 14 of aircraft 10 toward a rear end 16 of aircraft 10. As used herein, the term "fuselage" generally includes the entire body of aircraft 10, such as the tail of aircraft 10 and the outer surface or skin of aircraft 10. A first of wings 22 extends laterally outward from a port side 24 of fuselage 20 relative to longitudinal direction 12, and a second wing 22 extends laterally outward from a starboard side 26 of fuselage 20 relative to longitudinal direction 12. For the exemplary embodiment shown, each wing 22 includes one or more leading edge flaps 28 and one or more trailing edge flaps 30. Aircraft 10 also includes a vertical stabilizer 32 having rudder flaps 34 for yaw control and a pair of horizontal stabilizers 36 each having lift flaps 38 for pitch control. Fuselage 20 additionally includes an outer surface 40.

[0022] As shown, each stabilizer extends substantially in a single plane between the root portion and the tip portion. Figure 1 and 2 As shown, the vertical stabilizer 32 defines a root portion 60 and a tip portion 62 separated along a vertical direction V. Furthermore, the vertical stabilizer 32 extends along the longitudinal direction 12 between a leading edge 64 and a trailing edge 66. As shown, the vertical stabilizer 32 is mounted to the fuselage 20 at the root portion 60 and extends substantially along the vertical direction V to the tip portion 62. In this manner, a connection line 68 is defined at the intersection of the vertical stabilizer 32 and the fuselage 20. More specifically, the connection line 68 extends between the leading edge 64 and the trailing edge 66 of the vertical stabilizer 32. However, it should be understood that in other exemplary embodiments of the present disclosure, the aircraft 10 may additionally or alternatively include any other suitable stabilizer configuration that may or may not extend directly along the vertical direction V or the horizontal / lateral direction L. Furthermore, the alternative stabilizer may be of any suitable shape, size, configuration, or orientation while remaining within the scope of the present subject matter.

[0023] Figure 1 and 2 The exemplary aircraft 10 also includes a propulsion system. The exemplary propulsion system includes a plurality of aircraft engines, at least one of which is mounted to each of the pair of wings 22. Specifically, the plurality of aircraft engines includes a first aircraft engine 42 mounted to a first wing of the pair of wings 22 and a second aircraft engine 44 mounted to a second wing of the pair of wings 22. In at least some exemplary embodiments, the aircraft engines 42, 44 can be configured as turbofan jet engines suspended below the wings 22 in an underwing configuration. For example, in at least some exemplary embodiments, the first and / or second aircraft engines 42, 44 can be configured as turbofan jet engines as described below with reference to FIG. Figure 3The exemplary turbofan jet engine 100 described above is configured in substantially the same manner. However, alternatively, 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 42, 44 may alternatively be configured as turbojets, turboshafts, turboprops, etc.

[0024] In addition, the propulsion system includes an aft engine 200 mounted to the fuselage 20 of the aircraft 10 near the rear end 16 of the aircraft 10, or more specifically, at a location behind the wing 22 and the aircraft engines 42, 44. The exemplary aft engine 200 is mounted to the fuselage 20 of the aircraft 10 such that the mean line 18 extends therethrough. The aft engine 200 is generally configured to ingest and consume air forming a boundary layer on the fuselage 20, as will be referenced below. Figures 4 to 7 Discuss in more detail.

[0025] Specific reference Figure 2 , the aircraft 10 additionally includes landing gear, such as wheels 46, extending from the underside of the fuselage 20 and from the undersides of the wings 22. The fuselage 20 is designed to allow the aircraft 10 to take off and / or land at a takeoff angle 48 with the ground without the rear end 16 scraping the ground. More specifically, the takeoff angle 48 may be defined as the angle between the ground (parallel to the longitudinal direction 12) and a takeoff plane 50. As will be discussed below, despite the addition of the rear engine 200 near the rear end 16 of the aircraft 10, the exemplary fuselage 20 and rear engine 200 described herein are designed to allow the aircraft 10 to maintain a desired takeoff angle 48. Notably, for the depicted embodiment, the longitudinal direction 12 of the aircraft 10 is parallel to the ground when the aircraft 10 is on the ground. Thus, as shown, the maximum takeoff angle 48 may alternatively be defined by the longitudinal direction 12 of the aircraft 10 (at Figure 2 Angle 48' is shown in FIG.

[0026] Now refer to Figure 3 , provides a schematic cross-sectional view of an exemplary aircraft engine. Specifically, for the depicted embodiment, the aircraft engine is configured as a high-bypass turbofan jet engine, referred to herein as "turbofan engine 100." As described above, Figure 1 and Figure 2 One or both of the first and / or second aircraft engines 42, 44 of the exemplary aircraft 10 described herein may be operated in conjunction with Figure 3 However, alternatively, in other exemplary embodiments, one or both of the aircraft engines 42, 44 may be configured as any other suitable engine, such as a turboshaft engine, a turboprop engine, a turbojet engine, etc.

[0027] like Figure 3 As shown, turbofan engine 100 defines an axial direction A1 (extending parallel to a longitudinal centerline 102 for reference) and a radial direction R1 . Generally, turbofan 10 includes a fan section 104 and a core turbine engine 106 disposed downstream of fan section 104 .

[0028] The exemplary core turbine engine 106 shown generally includes a substantially tubular casing 108 defining an annular inlet 110. Casing 108 encloses, in serial flow relationship, a compressor section including a supercharger or low-pressure (LP) compressor 112 and a high-pressure (HP) compressor 114; a combustion section 116; a turbine section including a high-pressure (HP) turbine 118 and a low-pressure (LP) turbine 120; and an exhaust nozzle section 122. A high-pressure (HP) shaft or spool 124 drivingly connects the HP turbine 118 to the HP compressor 114. A low-pressure (LP) shaft or spool 126 drivingly connects the LP turbine 120 to the LP compressor 112. The compressor section, combustion section 116, turbine section, and nozzle section 122 together define a core air flow path.

[0029] In the depicted embodiment, fan section 104 includes a variable-pitch fan 128 having a plurality of fan blades 130 coupled to a disk 132 in a spaced-apart manner. As shown, fan blades 130 generally extend outward from disk 132 in a radial direction R1 and define a fan diameter D. Each fan blade 130 is rotatable relative to disk 132 about a pitch axis P by virtue of fan blades 130 being operably coupled to a suitable actuation member 134 configured to collectively and uniformly vary the pitch of fan blades 130. According to alternative embodiments, fan blades 130 may alternatively have a fixed pitch. Fan blades 130, disk 132, and actuation member 134 are collectively rotatable about longitudinal direction I2 via a power gearbox 136 that spans LP shaft 126. Power gearbox 136 includes a plurality of gears for adjusting the rotational speed of fan 128 relative to LP shaft 126 to a more efficient rotational fan speed. It should be understood that power gearbox 136 is optional and may not be used in some embodiments.

[0030] Still refer to Figure 3In the exemplary embodiment, the disk 132 is covered by a rotatable forward hub 138 that is aerodynamically shaped to facilitate airflow through the plurality of fan blades 130. In addition, the exemplary fan section 104 includes an annular fan casing or outer nacelle 140 that circumferentially surrounds the fan 128 and / or at least a portion of the core turbine engine 106. It should be understood that the nacelle 140 can be configured to be supported relative to the core turbine engine 106 by a plurality of circumferentially spaced outlet guide vanes 142. Furthermore, a downstream portion 144 of the nacelle 140 can extend over an outer portion of the core turbine engine 106 to define a bypass airflow passage 146 therebetween.

[0031] However, it should be understood that Figure 3 The exemplary turbofan engine 100 depicted in FIG. 1 is merely exemplary, and in other exemplary embodiments, turbofan engine 100 may have any other suitable configuration including, for example, any suitable number of shafts or spools, compressors, and / or turbines.

[0032] Now also refer to Figure 4 , provides Figure 1 and Figure 2 FIG2 is a close-up schematic cross-sectional view of an exemplary aft engine 200. As discussed, the exemplary aft engine 200 is mounted to the fuselage 20 near the aft end 16 of the aircraft 10. The illustrated aft engine 200 defines an axial direction A2 extending along a longitudinal centerline axis 220, a radial direction R2, and a circumferential direction C2 (not shown), with the longitudinal centerline axis 220 extending therethrough for reference.

[0033] Additionally, for the depicted embodiment, the aft engine 200 is configured as a boundary layer ingestion engine that is configured to ingest and consume air that forms a boundary layer on the fuselage 20 of the aircraft 10. The aft engine 200 includes a fan 222 rotatable about a central axis 220, a nacelle 224 extending around a portion of the fan 222, and one or more structural members 226 extending between the nacelle 224 and the fuselage 20 of the aircraft 10. The fan 222 includes a plurality of fan blades 228 that are generally spaced apart in a circumferential direction C2. Additionally, the nacelle 224 extends around and surrounds the plurality of fan blades 228 and a portion of the fuselage 20. Specifically, when the fan 222 is rotated about a central axis 220, the aft engine 200 may be configured to ingest and consume air that forms a boundary layer on the fuselage 20 of the aircraft 10. Figure 4 When the rear engine 200 is installed on the aircraft 10 , the nacelle 224 extends around at least a portion of the fuselage 20 of the aircraft 10 .

[0034] like Figure 4As also depicted in FIG, fan 222 further includes a fan shaft 230 to which a plurality of fan blades 228 are attached. Although not shown, fan shaft 230 may be rotatably supported by one or more bearings located in front of the plurality of fan blades 228, and optionally, may be rotatably supported by one or more bearings located behind the plurality of fan blades 228. Such bearings may be any suitable combination of roller bearings, ball bearings, thrust bearings, and the like.

[0035] In certain exemplary embodiments, plurality of fan blades 228 may be fixedly attached to fan shaft 230, or alternatively, plurality of fan blades 228 may be rotatably attached to fan shaft 230. For example, plurality of fan blades 228 may be attached to fan shaft 230 such that the pitch of each of plurality of fan blades 228 may be changed, for example, in unison, by a pitch changing mechanism (not shown).

[0036] The fan shaft 230 is mechanically coupled to a power source 232, which is located at least partially within the fuselage 20 of the aircraft 10. For the depicted embodiment, the fan shaft 230 is mechanically coupled to the power source 232 via a gearbox 234. The gearbox 234 can be configured to change the rotational speed of the power source 232, or more specifically, the rotational speed of a shaft 236 of the power source 232, so that the fan 222 of the rear engine 200 rotates at a desired rotational speed. The gearbox 234 can be a fixed ratio gearbox, or alternatively, the gearbox 234 can define a variable gear ratio.

[0037] The power source 232 can be any suitable power source. For example, in certain exemplary embodiments, the power source 232 can be an electrical source (e.g., the aft engine 200 can be configured as part of a gas-electric propulsion system with the first and / or second aircraft engines 42, 44). However, in other exemplary embodiments, the power source 232 can alternatively be configured as a dedicated gas engine, such as a gas turbine engine. Furthermore, in certain exemplary embodiments, the power source 232 can be positioned, for example, within the fuselage 20 of the aircraft 10 or within the aft engine 200 at any other suitable location. For example, in certain exemplary embodiments, the power source 232 can be configured as a gas turbine engine positioned at least partially within the aft engine 200.

[0038] Still refer to Figure 4, one or more structural members 226 extend between the nacelle 224 and the fuselage 20 of the aircraft 10 at a position forward of the plurality of fan blades 228. The one or more structural members 226 of the illustrated embodiment extend substantially along the radial direction R2 between the nacelle 224 and the fuselage 20 of the aircraft 10 for mounting the rear engine 200 to the fuselage 20 of the aircraft 10. However, it should also be understood that in other exemplary embodiments, the one or more structural members 226 may alternatively extend substantially along the axial direction A2, or in any other suitable direction between the axial direction A2 and the radial direction R2. It should be understood that as used herein, approximate terms such as "approximately," "substantially," or "approximately" refer to within a 10% error range.

[0039] The depicted one or more structural members 226 are configured as inlet guide vanes for the fan 222, such that the one or more structural members 226 are shaped and oriented to direct and condition air flow into the aft engine 200 to increase the efficiency of the aft engine 200. In certain exemplary embodiments, the one or more structural members 226 may be configured as fixed inlet guide vanes extending between the nacelle 224 and the fuselage 20 of the aircraft 10, or alternatively, the one or more structural members 226 may be configured as variable inlet guide vanes.

[0040] Furthermore, the aft engine 200 includes one or more outlet guide vanes 238 and a tail cone 240. For the illustrated embodiment, the one or more outlet guide vanes 238 extend between the nacelle 224 and the tail cone 240, for example, to increase the strength and rigidity of the aft engine 200 and to direct airflow through the aft engine 200. The outlet guide vanes 238 may be evenly spaced along the circumferential direction C2, or may have any other suitable spacing. Furthermore, the outlet guide vanes 238 may be fixed outlet guide vanes, or alternatively, variable outlet guide vanes.

[0041] Behind the plurality of fan blades 228, and for the depicted embodiment, behind one or more outlet guide vanes 238, the aft engine 200 further defines a nozzle 242 between the nacelle 224 and a tail cone 240. The nozzle 242 may be configured to generate a certain amount of thrust from the air flowing therethrough, and the tail cone 240 may be shaped to minimize drag on the aft engine 200. However, in other embodiments, the tail cone 240 may have any other shape and may, for example, terminate forward of the aft end of the nacelle 224 such that the tail cone 240 is surrounded by the nacelle 224 at the aft end. Additionally, in other embodiments, the aft engine 200 may not be configured to generate any measurable amount of thrust, but may instead be configured to ingest air from the boundary layer of air over the fuselage 20 of the aircraft 10 and add energy / accelerate this air to reduce the overall drag on the aircraft 10 (thereby increasing the net thrust of the aircraft 10).

[0042] Still refer to Figure 4 The aft engine 200, or more specifically the nacelle 224, defines an inlet 244 at a forward end 246 of the nacelle 224. The inlet 244 is defined by the nacelle 224 and the fuselage 20, i.e., between the nacelle 224 and the fuselage 20. As described above, the nacelle 224 of the aft engine 200 extends around and encloses the plurality of fan blades 228 of the fan 222 of the aft engine 200. For the depicted embodiment, the nacelle 224 also extends at least partially around the central axis 220 of the aft engine 200 and at least partially around the meanline 18 of the aircraft 10. Specifically, for the depicted embodiment, the nacelle 224 extends approximately three hundred and sixty degrees (360°) around the central axis 220 of the aft engine 200 and approximately three hundred and sixty degrees (360°) around the meanline 18 of the aircraft 10.

[0043] Notably, by positioning the rear engine 200 such that the nacelle 224 of the rear engine 200 extends at least partially around the fuselage 20 near the rear end 16 of the aircraft 10, for example, the bottom portion 248 of the nacelle 224 may not interfere with the takeoff angle 48 of the aircraft 10 (see FIG. Figure 2 For example, as shown, the nacelle 224 of the rear engine 200 includes at least a portion located inboard of the takeoff plane 50 defined by the fuselage 20 (see FIG. Figure 2 ). Particularly for the illustrated embodiment, the entire bottom portion 248 of the nacelle 224 is positioned in line with or inward of the takeoff plane 50 of the fuselage 20. For at least some prior art aircraft, the takeoff plane 50 of the fuselage 20 represents a conventional shape for the bottom portion of the fuselage at the rear end of the aircraft.

[0044] Now refer to Figures 4 to 7, the injection assembly 300 configured to provide improved boundary layer uptake will be described in greater detail. More specifically, Figures 4 to 7 Injection assemblies 300 having various configurations are provided as part of the aircraft 10. Although the various injection assemblies 300 are different, similar reference numerals will be used to describe them. Furthermore, while described with reference to the exemplary aircraft 10, it should be understood that the injection assembly 300 may be defined within the fuselage of any suitable aircraft to provide supplemental airflow to displace and / or redirect the primary boundary layer airflow over the fuselage. Depending on the specific application, the injection assembly 300 may also be modified as needed to increase the intake of boundary layer airflow into the rear engine 200.

[0045] Specific reference Figure 4 According to an exemplary embodiment, the air injection assembly 300 includes an injection port 302 defined on a surface of the fuselage 20. More specifically, according to the illustrated embodiment, the injection port 302 is positioned on the top side 202 of the fuselage 20 at an upstream position of the rear engine 200, and more specifically upstream of the inlet 244. The injection port 302 is fluidly coupled to a fluid passage 304 extending through the fuselage 20. According to Figure 4 In the exemplary embodiment of FIG. 3 , fluid channel 304 extends between inlet port 306 and injection port 302 . In this manner, fluid channel 304 receives a supplemental flow of gas (indicated by arrow 308 ) via inlet port 306 and supplies it to injection port 302 .

[0046] As described below, during operation of the aircraft 10, the primary airflow (indicated by arrow 310) may have a non-uniform velocity distribution along the circumferential direction C2 of the engine 200. Specifically, the primary airflow 310 on the top side of the fuselage 20 may have greater momentum than the primary airflow 310 on the bottom side of the fuselage 20. The inlet 306 is positioned to receive the supplemental airflow 308 from a location external to the fuselage 20 of the aircraft 10, and the injection port 302 is positioned to discharge the supplemental airflow 308 in a manner that at least partially moves and / or redirects the high-momentum primary airflow 310, for example, by pushing it upward in the vertical direction V. Figure 4 In the illustrated embodiment, the inlet 306 is defined in the underside 204 of the fuselage 20. Furthermore, the inlet 306 is positioned upstream of the injection port 302 along the longitudinal direction 12. The injection port 302 and the fluid passage 304 can be positioned and oriented so that the supplemental airflow 308 causes the higher velocity primary airflow 310 to bypass the fan inlet 244. In this manner, the low momentum supplemental airflow 308 enters the fan inlet 244 and results in a more uniform velocity distribution in the circumferential direction C2 about the aft engine 200.

[0047] For example, still referring to Figure 4, the top side 202 of the fuselage 202 defines a line, path, or trajectory for the primary airflow 310 (as indicated by the first reference line 312). Similarly, the fluid passage 304 and the injection port 302 define a direction for the supplemental airflow 308 (as indicated by the second reference line 314). According to the illustrated embodiment, the first reference line 312 (corresponding to the primary airflow 310) and the second reference line 314 (corresponding to the supplemental airflow 308) define an injection angle 316. The injection angle 316 can be designed to ensure that the supplemental airflow 308 exits the injection port 302 along a trajectory that redirects the primary airflow 310 as desired. For example, the injection angle 316 can be between five degrees and eighty-five degrees. According to other embodiments, the injection angle 316 is less than forty-five degrees. The injection angle 316 can be selected so that the relatively low velocity supplemental airflow 308 at least partially displaces or redirects the relatively high velocity primary airflow 310 so that the resulting velocity profile entering the inlet 244 is substantially constant along the circumferential direction C2. As Figure 5 As shown, injection port 302 is configured such that injection angle 316 is partially aligned with primary airflow 310 .

[0048] Still refer to Figure 4 , the injection assembly 300 may also include various features for controlling the flow of supplemental air 308. For example, the exemplary injection assembly 300 shown includes a booster fan 320. The booster fan 320 is in fluid communication with the fluid passage 304 to increase the pressure or velocity of the supplemental air flow 308. Additionally, the injection assembly 300 includes a valve 322 for controlling the flow of supplemental air 308 through the fluid passage 304. For the depicted embodiment, the booster fan 320 is an electric fan positioned within the fluid passage 304, and the valve 322 is a variable throughput valve configured to control the flow rate of the supplemental air flow 308.

[0049] In addition, injection assembly 300 includes one or more sensors, including, for example, an airflow duct sensor 324 positioned in or adjacent to fluid passage 304, for determining one or more of a pressure or a velocity of supplemental airflow 308 passing through fluid passage 304. It is noted that aircraft 10 also includes additional sensors for measuring parameters of aircraft 10, which may also be used to control injection assembly 300. For example, aircraft 10 includes sensor 326 for measuring at least one of a pressure or a velocity of airflow entering fan inlet 244.

[0050] Furthermore, for the depicted embodiment, the injection assembly 300 further includes a controller 330, which in certain exemplary embodiments may be configured as part of a main aircraft controller and operatively connected to the booster fan 320, valve 322, and sensors 324, 326 via one or more wired or wireless communication buses. The controller 330 may be configured to operate the booster fan 320 and / or valve 322 at least in part in response to the pressure and / or speed measured by the sensors 324, 326.

[0051] Controller 330 may be any suitable controller. For example, controller 330 may include one or more computing devices having one or more processors and one or more memory devices for storing information accessible to the one or more processors. Controller 330 may also include a communication interface operably connected to one or more wired or wireless communication buses, allowing controller 330 to communicate with various other components of aircraft 10.

[0052] However, it should be understood that Figure 4 The exemplary embodiments depicted in FIG. 3 are provided as examples only. In other exemplary embodiments, the aircraft 10 and the injection assembly 300 may alternatively be configured in any other suitable manner. For example, referring now to FIG. Figure 5 and Figure 6 , depicts a cross-sectional side view of an alternative injection assembly 300 defined in the fuselage 20 of the aircraft 10 , according to other exemplary embodiments of the present disclosure. Figure 5 and Figure 6 The exemplary aircraft 10 may be operated with Figure 4 The exemplary aircraft 10 is configured in substantially the same manner and, therefore, like reference numerals refer to like or similar parts.

[0053] refer to Figure 5 , the aircraft 10 includes a fluid passage 304 extending between an inlet 306 positioned to receive airflow from a location external to the fuselage 20, and an injection port 302 positioned to discharge a supplemental airflow 308 to at least partially displace or redirect a primary airflow 310 upstream of the aft engine 200. However, for the depicted embodiment, the inlet 306 of the fluid passage 304 includes a plurality of inlets 306 defined by the fuselage 12 of the aircraft 10 at the bottom side 204 of the fuselage 20 of the aircraft 10. Additionally, for the depicted embodiment, the injection port 302 of the fluid passage 304 includes a plurality of injection ports 302. The plurality of injection ports 302 are arranged along the top side 202 of the fuselage 20 upstream of the aft engine 200.

[0054] The plurality of inlets 306 and injection ports 302 can be arranged along the fuselage 20 in any suitable manner to achieve the desired supplemental airflow 308 for a particular application. For example, according to an exemplary embodiment, the inlets 306 and injection ports 302 can be spaced apart along the longitudinal direction 12 (as shown) and / or along the circumferential direction C2 (not shown). The inlets 306 and injection ports 302 can be arranged linearly, in a zigzag pattern, or in any other suitable pattern. Other patterns and configurations are possible and within the scope of the present subject matter.

[0055] Additionally, each of the inlets 306 and injection ports 302 can be substantially evenly spaced and of substantially the same size, or alternatively, can vary in spacing and / or size to more efficiently supply the supplemental airflow 308 as desired. For example, the inlets 306 and injection ports 302 can be circular, oval, square, rectangular, or any other suitable shape or combination of shapes. Similarly, the fluid passage 304 can be arranged along a straight line between the inlets 306 and injection ports 302, or can be arranged along any suitable path through the fuselage 20 to improve the velocity, injection angle, or flow rate of the supplemental airflow 308. Furthermore, it should be understood that in other exemplary embodiments, the inlets 306 can include any suitable number of inlets 306, and similarly, the injection ports 302 can include any suitable number of injection ports 302.

[0056] for Figure 5 In the embodiment depicted in FIG, the plurality of inlet ports 306 converge into a single body portion 332 of the fluid channel 304 before subsequently separating into the plurality of injection ports 302. However, in other embodiments, for example Figure 6 , fluid passage 304 may include a plurality of fluid passages 304 extending from a respective inlet 306 to a respective injection port 302. Each respective inlet 306 may be defined by fuselage 20 of aircraft 10 at underside 204 of fuselage 20, or alternatively, may be defined at any other suitable location.

[0057] refer to Figure 7 According to some exemplary embodiments, the fluid passage 304 may receive the supplemental airflow 308 from a location other than the surface of the fuselage 20. For example, as shown, the fluid passage 304 is configured to draw bleed air from a compressor (e.g., the LP compressor 112) of a gas turbine engine 100 (e.g., the first or second aircraft engine 42, 44). Furthermore, one or more pressure regulators may be used to adjust the pressure of the supplemental airflow 308 received from the LP compressor 112 to achieve a desired velocity, flow rate, etc. of the supplemental airflow 308. Other embodiments may include receiving high-pressure air from the fan 222 or downstream of the rear engine 200.

[0058] The injection assembly 300, as described above, can be used to more effectively distribute the boundary layer airflow ingested by the aft engine 200. The shape, size, and configuration of the injection port and fluid passageway can be optimized to displace the boundary layer airflow to maximize the ingestion of low-speed air through the aft engine 20, thereby providing the aft engine 200 with a more uniform velocity distribution circumferentially around the fuselage 20 and the fan inlet 244, and improving the propulsive efficiency of the aircraft 10. Additionally, the velocity of the boundary layer air flowing into the aft engine 200 can be similar from the upper half to the lower half, thereby improving propulsive efficiency while reducing vibration, noise, and wear on the plurality of fan blades 228. The source of supplemental airflow 308 can be drawn from any suitable location within or around the aircraft 10 and can be operatively coupled to the injection port 302 in any suitable manner to achieve optimal displacement of the boundary layer airflow. However, it is noteworthy that drawing the supplemental airflow 308 from the lower half of the fuselage 20 results in a lower pressure along the lower half, thereby reducing the pressure differential between the upper and lower halves and reducing vortex distortion.

[0059] This written description uses examples to disclose the invention, including the best mode, and also to enable those skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims and may include other examples that occur to those skilled in the art. If such other examples include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims, such other examples are intended to be within the scope of the claims.

Claims

1. An aircraft, characterized in that: The aircraft defines a longitudinal direction, a vertical direction, and a lateral direction, and the aircraft comprises: a fuselage extending along the longitudinal direction between a front end and a rear end; a boundary layer intake fan mounted to the fuselage at an aft end of the fuselage, the boundary layer intake fan defining a centerline and including a plurality of fan blades rotatable about the centerline; and Air injection assembly, including: a plurality of injection ports defined on the top surface of the fuselage at a location upstream of the boundary layer ingestion fan; a fluid passage extending through at least a portion of the fuselage and in fluid communication with the plurality of injection ports for directing a flow of supplemental air through the plurality of injection ports; and at least one inlet port defined on a bottom surface of the fuselage, wherein the fluid passage extends between the at least one inlet port and the plurality of injection ports, and Wherein, the injection assembly further includes a booster fan, which is in fluid communication with the fluid channel and is used to increase the pressure of the supplementary airflow.

2. The aircraft according to claim 1, characterized in that in, The at least one inlet further includes a plurality of inlets defined on the bottom surface of the fuselage, and wherein the fluid passage extends between the plurality of inlets and the plurality of injection ports.

3. The aircraft according to claim 1, characterized in that in, The boost fan is disposed in the fluid channel.

4. The aircraft according to claim 1, characterized in that further comprising a wing-mounted gas turbine engine including a compressor, the wing-mounted gas turbine engine being disposed longitudinally forward of the boundary layer ingestion fan, and Wherein the fluid passage is configured for extracting bleed air from the compressor of the wing-mounted gas turbine engine of the aircraft.

5. The aircraft according to claim 1, characterized in that wherein the fluid channel comprises a plurality of fluid channels, wherein each of the plurality of fluid channels comprises at least one valve disposed therein, and Wherein, the at least one valve comprises a variable flow valve.

6. The aircraft according to claim 1, characterized in that in, The surface of the fuselage defines a primary airflow reference line, and the fluid channel and the plurality of injection ports define a secondary reference line, such that the supplemental airflow is discharged along the secondary reference line, wherein the primary airflow reference line and the auxiliary reference line define an injection angle between five degrees and 85 degrees, and wherein the primary airflow merges with the supplemental airflow behind the primary airflow reference line at the top surface of the fuselage upstream of the boundary layer ingestion fan.

7. The aircraft according to claim 6, characterized in that in, The injection angle is less than 45 degrees, wherein the velocity of the primary airflow is higher than the velocity of the supplementary airflow, and wherein the supplemental airflow causes at least a portion of the primary airflow to bypass the boundary layer ingestion fan.

8. The aircraft according to claim 1, characterized in that in, The plurality of injection ports are spaced apart along a circumferential direction.

9. The aircraft according to claim 1, characterized in that wherein the plurality of injection ports are substantially unevenly spaced, and Wherein, the fluid channel includes at least one curved portion.

10. The aircraft according to claim 9, characterized in that in, The plurality of injection ports have different sizes, and Wherein, the fluid channel includes at least two curved parts.

11. The aircraft according to claim 1, characterized in that in, The fluid channel includes at least two curved portions.

12. An air injection system for an aircraft, characterized in that: include: a fuselage extending in a longitudinal direction between a forward end and a rearward end, and a boundary layer intake fan mounted to the fuselage at a rearward end of the fuselage; a plurality of injection ports defined on a surface of the fuselage at a location upstream of the boundary layer ingestion fan; a fluid passage extending from at least one inlet provided on the bottom surface of the body through the body to the plurality of injection ports provided on the top surface of the body, wherein the plurality of injection ports have different sizes, and The fluid channel includes at least two curved portions.

13. The system according to claim 12, wherein: The invention further includes a booster fan in fluid communication with the fluid channel to increase the pressure of the supplementary air flow, the booster fan being controlled by at least one controller.

14. The system according to claim 12, wherein: in, The aircraft further includes a wing-mounted gas turbine engine disposed longitudinally forward of the boundary layer ingestion fan, the wing-mounted gas turbine engine including a compressor, and The fluid passage is configured to extract bleed air from a compressor of the gas turbine engine of the aircraft.

15. The system according to claim 12, wherein: in, The at least one inlet further includes a plurality of inlets defined on a bottom surface of the body, and the fluid passage extends between the plurality of inlets and the plurality of injection ports.

16. The system according to claim 12, wherein: in, The fluid channel includes a plurality of fluid channels, and Wherein, each cross section of the plurality of injection ports and the at least one inlet is at least one of an elliptical, square and rectangular shape.

17. The system according to claim 13, wherein: The surface of the fuselage defines a primary airflow reference line, and the fluid channel and the plurality of injection ports define a secondary reference line, such that the supplemental airflow is discharged along the secondary reference line, and The primary airflow reference line and the auxiliary reference line define an injection angle between five degrees and 85 degrees.

18. The system according to claim 17, wherein: in, The injection angle is less than 45 degrees.

19. The system according to claim 12, wherein: in, The multiple injection ports are substantially uniform in size.

20. An aircraft, the aircraft defining a longitudinal direction, a vertical direction and a lateral direction, characterized in that: The aircraft comprises: a fuselage extending along the longitudinal direction between a front end and a rear end; a boundary layer intake fan mounted to the fuselage at an aft end of the fuselage, the boundary layer intake fan defining a centerline and including a plurality of fan blades rotatable about the centerline; and An air injection assembly, the air injection assembly comprising: at least one injection port defined on the top surface of the fuselage at a location upstream of the boundary layer ingestion fan; a fluid passage extending through the fuselage and fluidly coupled upstream of the at least one injection port, the fluid passage further comprising: at least one valve disposed therein; at least one first sensor disposed therein; and A booster fan is disposed therein, wherein the booster fan is disposed upstream of the at least one valve.

21. The aircraft according to claim 20, characterized in that Further including: at least one second sensor disposed on a radially inward portion of a nacelle of the boundary layer ingestion fan, the at least one second sensor configured to measure at least one of a pressure and a velocity of an airflow entering the boundary layer ingestion fan; and at least one controller communicatively coupled to the at least one first sensor and the at least one second sensor, wherein the at least one controller controls at least one of the booster fan and the at least one valve based at least in part on at least one signal from at least one of the at least one first sensor and the at least one second sensor, and Wherein, the at least one valve comprises a variable flow valve.

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