Abstract
An assembly for an aircraft propulsion system includes a propulsor rotor and a guide vane structure. The guide vane structure includes a plurality of guide vanes. The guide vane structure is disposed axially next to the propulsor rotor. The guide vanes include a first guide vane. The first guide vane includes a camber line, a leading edge, a trailing edge, an inner end, an outer end, at least one leading edge section and a trailing edge section. The first guide vane extends longitudinally along the camber line from the leading edge to the trailing edge. The leading edge section forms a respective portion of the leading edge and is arranged radially along the trailing edge section between the inner end and the outer end. The trailing edge section forms the trailing edge. The leading edge section is configured to translate longitudinally along the camber line.
Claims
1. An assembly for an aircraft propulsion system, comprising: a propulsor rotor configured to rotate about an axis; and a guide vane structure including a plurality of guide vanes arranged circumferentially about the axis, the guide vane structure disposed axially next to the propulsor rotor, and the plurality of guide vanes comprising a first guide vane; the first guide vane comprising a camber line, a leading edge, a trailing edge, an inner end, an outer end, at least one leading edge section and a trailing edge section, the first guide vane extending longitudinally along the camber line from the leading edge to the trailing edge, the leading edge section forming a respective portion of the leading edge and arranged radially along the trailing edge section between the inner end and the outer end, and the trailing edge section forming the trailing edge; and the at least one leading edge section configured to translate longitudinally along the camber line relative to the trailing edge section while maintaining longitudinal overlap between the leading edge section and the trailing edge section.
2. The assembly of claim 1, wherein a first of the at least one leading edge section is configured to translate longitudinally along the camber line independent of a second of the at least one leading edge section.
3. The assembly of claim 1, wherein a first of the at least one leading edge section is further configured to translate longitudinally along the camber line relative to a second of the at least one leading edge section.
4. The assembly of claim 1, wherein the at least one leading edge section include a first leading edge section forming a first portion of the leading edge, the first leading edge section configured to translate longitudinally along the camber line from a first section retracted position to a first section extended position; and a second leading edge section disposed radially outboard of the first leading edge section, the second leading edge section forming a second portion of the leading edge, and the second leading edge section configured to translate longitudinally along the camber line from a second section retracted position to a second section extended position.
5. The assembly of claim 4, wherein the first portion of the leading edge is aligned with the second portion of the leading edge when the first leading edge section is in the first section retracted position and the second leading edge section is in the second section retracted position.
6. The assembly of claim 4, wherein the first portion of the leading edge is aligned with the second portion of the leading edge when the first leading edge section is in the first section extended position and the second leading edge section is in the second section extended position.
7. The assembly of claim 4, wherein the first portion of the leading edge is offset from the second portion of the leading edge during at least one mode of operation.
8. The assembly of claim 4, wherein the first leading edge section is longitudinally translated farther than the second leading edge section during at least one mode of operation.
9. The assembly of claim 4, wherein the second leading edge section is longitudinally translated farther than the first leading edge section during at least one mode of operation.
10. The assembly of claim 1, wherein the trailing edge section is a stationary section of the first guide vane.
11. The assembly of claim 1, further comprising: an inner case structure; and an outer case structure axially overlapping and circumscribing the inner case structure; the trailing edge section structurally tying and configured to transfer loads between the inner case structure and the outer case structure.
12. The assembly of claim 1, wherein a first of the at least one leading edge section is configured to translate longitudinally along the camber line between a retracted position and an extended position; and the trailing edge section projects longitudinally into a channel in the first of the at least one leading edge section when the first of the at least one leading edge section is in the retracted position.
13. The assembly of claim 12, wherein the trailing edge section projects longitudinally into the channel in the first of the at least one leading edge section when the first of the at least one leading edge section is in the extended position.
14. The assembly of claim 1, wherein a first of the at least one leading edge section is configured to translate longitudinally along the camber line between a retracted position and an extended position; the first of the at least one leading edge section includes a base and a side fairing projecting longitudinally out from the base towards the trailing edge; and the side fairing longitudinally overlaps the trailing edge section when the first of the at least one leading edge section is in the retracted position and the extended position.
15. The assembly of claim 1, further comprising a turbine engine core configured to drive rotation of the propulsor rotor about the axis, the turbine engine core including a flowpath, a compressor section, a combustor section and a turbine section, the flowpath extending through the compressor section, the combustor section and the turbine section from an inlet into the flowpath to an exhaust from the flowpath.
16. The assembly of claim 1, wherein the guide vane structure is disposed downstream of the propulsor rotor.
17. The assembly of claim 1, further comprising: a case structure axially overlapping and circumscribing the propulsor rotor and the guide vane structure; the propulsor rotor comprising a fan rotor.
18. An assembly for an aircraft propulsion system, comprising: a fan rotor configured to rotate about an axis; an inner case structure axially next to the fan rotor; an outer case structure axially overlapping and circumscribing the fan rotor and the inner case structure; and a guide vane structure including a plurality of guide vanes arranged circumferentially about the axis and radially between the inner case structure and the outer case structure, the guide vane structure disposed axially next and downstream of the fan rotor, and the plurality of guide vanes comprising a first guide vane; the first guide vane comprising a camber line, a leading edge, a trailing edge, a leading edge section and a trailing edge section, the first guide vane extending longitudinally along the camber line from the leading edge to the trailing edge, the leading edge section at least partially forming the leading edge, and the trailing edge section forming the trailing edge; the leading edge section configured to translate longitudinally along the camber line relative to the trailing edge section; and the trailing edge section structurally tying and configured to transfer loads between the inner case structure and the outer case structure.
19. The assembly of claim 18, wherein the leading edge section is one of a plurality of leading edge sections arranged radially along the trailing edge section, and each of the plurality of leading edge sections is configured to translate longitudinally along the camber line relative to the trailing edge section.
20. An assembly for an aircraft propulsion system, comprising: a fan rotor configured to rotate about an axis; an inner case structure axially next to the fan rotor; an outer case structure axially overlapping and circumscribing the fan rotor and the inner case structure; and a guide vane structure including a plurality of guide vanes arranged circumferentially about the axis and radially between the inner case structure and the outer case structure, the guide vane structure disposed axially next and downstream of the fan rotor, and the plurality of guide vanes comprising a first guide vane; the first guide vane comprising a camber line, a leading edge, a trailing edge, a leading edge section and a trailing edge section, the first guide vane extending longitudinally along the camber line from the leading edge to the trailing edge, the leading edge section at least partially forming the leading edge, and the trailing edge section forming the trailing edge; and the leading edge section configured to translate longitudinally along the camber line relative to the trailing edge section from a retracted position to an extended position, the leading edge section including a base and a side fairing that projects longitudinally out from the base towards the trailing edge, and the side fairing longitudinally overlapping the trailing edge section when the leading edge section is in the retracted position and the extended position.
Description
BRIEF DESCRIPTION OF THE DRAWINGS
[0029] FIG. 1 is a partial schematic sectional illustration of an aircraft propulsion system.
[0030] FIG. 2 is a schematic sectional illustration of a portion of the aircraft propulsion system at a guide vane structure.
[0031] FIG. 3 is a cross-sectional illustration of a telescopic, structural guide vane in a retracted position.
[0032] FIG. 4 is a cross-sectional illustration of the telescopic, structural guide vane in an extended position.
[0033] FIGS. 5-8 are schematic sectional illustrations of a portion of the aircraft propulsion system with a telescopic, structural guide vane in various arrangements.
[0034] FIGS. 9-11 are schematic sectional illustrations of a portion of the aircraft propulsion system with various telescopic, structural guide vane arrangements.
DETAILED DESCRIPTION
[0035] FIG. 1 illustrates a propulsion system 20 for an aircraft. The aircraft may be an airplane, a drone (e.g., an unmanned aerial vehicle (UAV)) or any other manned or unmanned aerial vehicle or system. The aircraft propulsion system 20 of FIG. 1 includes a propulsor section 22 and a powerplant 24 configured to mechanically drive operation of the propulsor section 22. For ease of description, the powerplant 24 of FIG. 1 is described below as a core 26 (e.g., a gas generator) of a gas turbine engine 28 such as a turbofan engine, and the propulsor section 22 is described below as a ducted propulsor section such as a fan section of the turbofan engine. The present disclosure, however, is not limited to such an exemplary propulsion system configuration.
[0036] The aircraft propulsion system 20 of FIG. 1 extends axially along an axis 30 between an axial forward, upstream end 32 of the aircraft propulsion system 20 and an axial aft, downstream end 34 of the aircraft propulsion system 20. Briefly, the axis 30 may be a centerline axis of the aircraft propulsion system 20, the turbine engine 28 and/or one or more of its members. The axis 30 may also or alternatively be a rotational axis for one or more members of the turbine engine 28.
[0037] The turbine engine 28 of FIG. 1 includes the propulsor section 22 (e.g., the fan section), a compressor section 36, a combustor section 37 and a turbine section 38. The compressor section 36 of FIG. 1 includes a low pressure compressor (LPC) section 36A and a high pressure compressor (HPC) section 36B. The turbine section 38 of FIG. 1 includes a high pressure turbine (HPT) section 38A and a low pressure turbine (LPT) section 38B.
[0038] The engine sections 22 and 36A-38B may be arranged sequentially along the axis 30 within a stationary engine housing 40 for the turbine engine 28. The propulsor section 22 includes a bladed propulsor rotor 42; e.g., a fan rotor. The LPC section 36A includes a bladed low pressure compressor (LPC) rotor 43. The HPC section 36B includes a bladed high pressure compressor (HPC) rotor 44. The HPT section 38A includes a bladed high pressure turbine (HPT) rotor 45. The LPT section 38B includes a bladed low pressure turbine (LPT) rotor 46. These propulsion engine rotors 42-46 are housed within the engine housing 40. The engine housing 40 of FIG. 1, for example, includes an inner housing structure 48 and an outer housing structure 50. Here, at least (or only) the LPC section 36A, the HPC section 36B, the combustor section 37, the HPT section 38A and the LPT section 38B collectively form the engine core 26.
[0039] The inner housing structure 48 of FIG. 1 includes an inner case structure 52 (e.g., a core case) for the turbine engine 28, and an inner nacelle structure 54 (sometimes referred to as an inner fixed structure (IFS)). The inner case structure 52 is disposed radially outboard of, extends axially along and may circumscribe one or more or all of the engine sections 36A-38B and their respective engine rotors 43-46. The inner case structure 52 may thereby house and provide a support structure for the respective engine sections 36A-38B and their respective engine rotors 43-46. The inner nacelle structure 54 is configured to provide an aerodynamic cover over the engine core 26 and its inner case structure 52. The inner housing structure 48 and its members 52 and 54 may also form a radial inner peripheral boundary of a (e.g., annular) bypass flowpath 56 within the aircraft propulsion system 20.
[0040] The outer housing structure 50 of FIG. 1 includes an outer case structure 58 (e.g., a fan case) for the turbine engine 28, and an outer nacelle structure 60. The outer case structure 58 is disposed radially outboard of, extends axially along and may circumscribe the propulsor section 22 and its propulsor rotor 42. The outer case structure 58 may thereby house and provide a containment structure for the propulsor section 22 and its propulsor rotor 42. The outer case structure 58 is also disposed radially outboard of, extends axially along and may circumscribe the inner housing structure 48 and its inner case structure 52. The outer nacelle structure 60 is configured to provide an aerodynamic cover over the outer case structure 58. The outer housing structure 50 and its members 58 and 60 may also form a radial outer peripheral boundary of the bypass flowpath 56.
[0041] The HPC rotor 44 is coupled to and rotatable with the HPT rotor 45. The HPC rotor 44 of FIG. 1, for example, is connected to the HPT rotor 45 through a high speed shaft 62. At least (or only) the HPC rotor 44, the HPT rotor 45 and the high speed shaft 62 collectively form a high speed rotating assembly 64; e.g., a high speed spool of the engine core 26. This high speed rotating assembly 64 of FIG. 1 and its members 44, 45 and 62 are rotatable about the axis 30.
[0042] The LPC rotor 43 is coupled to and rotatable with the LPT rotor 46. The LPC rotor 43 of FIG. 1, for example, is connected to the LPT rotor 46 through a low speed shaft 66. At least (or only) the LPC rotor 43, the LPT rotor 46 and the low speed shaft 66 collectively form a low speed rotating assembly 68; e.g., a low speed spool of the engine core 26. This low speed rotating assembly 68 is further coupled to the propulsor rotor 42 through a drivetrain 70. This drivetrain 70 may be configured as a geared drivetrain, where a geartrain 72 (e.g., a transmission, a speed change device, an epicyclic geartrain, etc.) is disposed between and operatively couples the propulsor rotor 42 to the low speed rotating assembly 68 and its LPT rotor 46. With this arrangement, the propulsor rotor 42 may rotate at a different (e.g., slower) rotational speed than the low speed rotating assembly 68 and its LPT rotor 46. Alternatively, the drivetrain 70 may be configured as a direct drive drivetrain, where the geartrain 72 is omitted. With such an arrangement, the propulsor rotor 42 rotates at a common (the same) rotational speed as the low speed rotating assembly 68 and its LPT rotor 46. The low speed rotating assembly 68 of FIG. 1 and its members 43, 46 and 66 as well as the propulsor rotor 42 are rotatable about the axis 30.
[0043] During operation, ambient air from outside of the aircraft enters the aircraft propulsion system 20 and its turbine engine 28 through an airflow inlet 74. This air is directed across the propulsor section 22 and into a (e.g., annular) core flowpath 76 and the bypass flowpath 56. The core flowpath 76 of FIG. 1 extends sequentially through the LPC section 36A, the HPC section 36B, the combustor section 37, the HPT section 38A and the LPT section 38B from an airflow inlet 78 into the core flowpath 76 to a combustion products exhaust 80 out from the core flowpath 76 and the engine core 26. The air entering the core flowpath 76 may be referred to as core air. The bypass flowpath 56 extends through a bypass duct, which bypass flowpath 56 and bypass duct bypass (e.g., are disposed radially outboard of and extend along) the engine core 26 and the inner housing structure 48. The air within the bypass flowpath 56 may be referred to as bypass air.
[0044] The core air is compressed by the LPC rotor 43 and the HPC rotor 44 and is directed into a (e.g., annular) combustion chamber 82 of a (e.g., annular) combustor 84 in the combustor section 37. Fuel is injected into the combustion chamber 82 by one or more fuel injectors 86 and mixed with the compressed core air to provide a fuel-air mixture. This fuel-air mixture is ignited and combustion products thereof flow through and sequentially drive rotation of the HPT rotor 45 and the LPT rotor 46 about the axis 30. The rotation of the HPT rotor 45 and the LPT rotor 46 respectively drive rotation of the HPC rotor 44 and the LPC rotor 43 about the axis 30 and, thus, compression of the air received from the core inlet 78. The rotation of the LPT rotor 46 also drives rotation of the propulsor rotor 42. The rotation of the propulsor rotor 42 propels the bypass air through and out of the bypass flowpath 56. The propulsion of the bypass air may account for a majority of thrust generated by the turbine engine 28 of FIG. 1.
[0045] Referring to FIG. 2, the rotation of the propulsor rotor 42 may introduce swirl into the air propelled into the bypass flowpath 56. This phenomenon may, in part, be manifest due to a difference in loading of a propulsor blade 92 between a hub that is moving relatively slowly, and a tip of the propulsor blade 92 that is moving relatively fast through the air. The aircraft propulsion system 20 thereby includes an exit guide vane structure 88 disposed within the bypass flowpath 56 radially between the inner housing structure 48 and the outer housing structure 50. This guide vane structure 88 is configured to condition (e.g., straighten out) the bypass air propelled through the bypass flowpath 56 by the propulsor rotor 42, for example to remove or reduce the circumferential swirl and thereby enhance the forward thrust.
[0046] The propulsion system may incorporate a degree of twist and airfoil tailoring to optimize flow for typical engine cruise conditions. As such, a fixed fan exit guide vane may experience less than optimal performance under varied operating conditions as a function of altitude, speed and condition of flight. The guide vane structure 88 of present disclosure therefore includes a plurality of operationally adjustable telescopic, structural exit guide vanes 90 to facilitate tuning for varied operating conditions. This guide vane structure 88 and its guide vanes 90 are arranged axially next to (e.g., adjacent) the propulsor rotor 42 and its propulsor blades 92. The guide vane structure 88 and its guide vanes 90 of FIG. 2, for example, are arranged downstream of the propulsor rotor 42 and its propulsor blades 92, without (e.g., any) other elements axially therebetween to obstruct, turn and/or otherwise influence the air propelled by the propulsor rotor 42 to the guide vane structure 88, for example.
[0047] The guide vanes 90 are arranged and may be equispaced circumferentially about the axis 30 in an array; e.g., a circular array. Each of the guide vanes 90 extends spanwise along a span line of the respective guide vane 90 (e.g., radially relative to the axis 30) across the bypass flowpath 56 from an inner end 94 of the respective guide vane 90 to an outer end 96 of the respective guide vane 90. The vane inner end 94 of FIG. 2 is disposed radially adjacent the inner case structure 52. The vane outer end 96 of FIG. 2 is disposed radially adjacent the outer case structure 58.
[0048] Referring to FIGS. 3 and 4, each guide vane 90 extends longitudinally along a camber line 98 of the respective guide vane 90 from a leading edge 100 of the respective guide vane 90 to a trailing edge 102 of the respective guide vane 90. Each guide vane 90 extends laterally between and to opposing side exterior surfaces 104A and 104B (generally referred to as 104) of the respective guide vane 90. The first vane exterior surface 104A of FIGS. 3 and 4 is a convex, suction side surface of the respective guide vane 90. The second vane exterior surface 104B of FIGS. 3 and 4 is a concave, pressure side surface of the respective guide vane 90. Each of these vane exterior surfaces 104 extends longitudinally along the vane camber line 98 between and meet at the respective vane leading edge 100 and the respective vane trailing edge 102. Referring to FIG. 2, each vane element 100, 102, 104 (the vane exterior surface 104B not visible in FIG. 2) extends spanwise from the vane inner end 94/the inner case structure 52 to the vane outer end 96/the outer case structure 58.
[0049] Each guide vane 90 of FIG. 2 includes one or more leading edge sections 106A and 106B (generally referred to as 106) and a trailing edge section 108. The leading edge sections 106 of FIG. 2 are configured as movable slats for the respective guide vane 90. The trailing edge section 108 of FIG. 2 is configured as a stationary structure member of the respective guide vane 90.
[0050] The leading edge sections 106 of FIG. 2 are disposed at and may at least partially (or completely) collectively form the respective vane leading edge 100. The leading edge sections 106 of FIG. 2, for example, are arranged radially end-to-end along the trailing edge section 108 between the vane inner end 94/the inner case structure 52 and the vane outer end 96/the outer case structure 58. Each of these leading edge sections 106 extends longitudinally along the respective vane camber line 98 from a downstream end 110 of the respective leading edge section 106 to the respective vane leading edge 100. Each leading edge section 106 extends spanwise along the respective vane leading edge 100 from an inner end 112 of the respective leading edge section 106 to an outer end 114 of the respective leading edge section 106. The inner end 112 of the inner leading edge section 106A of FIG. 2 is disposed at (e.g., on, adjacent or proximate) the vane inner end 94/the inner case structure 52. The outer leading edge section 106B is disposed radially outboard of the inner leading edge section 106A, where the inner end 112 of the outer leading edge section 106B of FIG. 1 is radially adjacent the outer end 114 of the inner leading edge section 106A. The outer end 114 of the outer leading edge section 106B of FIG. 2 is disposed at the vane outer end 96/the outer case structure 58. Referring to FIGS. 3 and 4, each leading edge section 106 extends laterally between and to the vane exterior surfaces 104.
[0051] Each leading edge section 106 of FIGS. 3 and 4 includes a leading edge section base 116 and one or more leading edge section fairings 118A and 118B (generally referred to as 118); e.g., sidewalls, extensions, etc. The section base 116 is disposed at and forms respective portion of the vane leading edge 100. Each of the section fairings 118 is disposed at the downstream end 110 of the respective leading edge section 106. Each section fairing 118 of FIGS. 3 and 4, for example, projects out from the section base 116, longitudinally along the respective vane camber line 98, to the downstream end 110 of the respective leading edge section 106. The first side fairing 118A is disposed laterally at and partially forms the first vane exterior surface 104A. The second side fairing 118B is disposed laterally at and partially forms the second vane exterior surface 104B.
[0052] The section base 116, the first side fairing 118A and the second side fairing 118B of FIGS. 3 and 4 form a channel 120 in the respective leading edge section 106. This channel 120 projects longitudinally along the respective vane camber line 98 into the respective leading edge section 106 from the downstream end 110 of the respective leading edge section 106 to the section base 116. The channel 120 extends laterally within the respective leading edge section 106 between the first side fairing 118A and the second side fairing 118B. The channel 120 may extend spanwise within, into or through the respective leading edge section 106.
[0053] The trailing edge section 108 is disposed at and (e.g., completely) forms the respective vane trailing edge 102. The trailing edge section 108 of FIG. 2, for example, extends longitudinally along the respective vane camber line 98 from an upstream end 122 of the trailing edge section 108 to the respective vane trailing edge 102. The trailing edge section 108 extends spanwise along the respective vane trailing edge 102 from the vane inner end 94 to the vane outer end 96. This trailing edge section 108 is also fixedly connected to (e.g., formed integral with, mechanically fastened to, or otherwise attached to) the inner housing structure 48 and the outer housing structure 50. More particularly, the trailing edge section 108 of FIG. 2 is structurally tied to the inner case structure 52 and the outer case structure 58 such that the trailing edge sections 108 of the guide vane structure 88 collectively structurally tie and transfer loads between the inner case structure 52 and the outer case structure 58 during propulsion system operation. Referring to FIGS. 3 and 4, the trailing edge section 108 extends laterally between and to the vane exterior surfaces 104. With the foregoing arrangement, the trailing edge section 108 functionally combines a structural strut and a trailing edge portion of an airfoil into a single (e.g., monolithic) unit.
[0054] The trailing edge section 108 of FIGS. 3 and 4 projects longitudinally along the respective vane camber line 98 into the channel 120 of each leading edge section 106. The first side fairing 118A of FIGS. 3 and 4 is laterally adjacent a first side 124A of the trailing edge section 108, and longitudinally and spanwise overlaps the trailing edge section 108 and its first side 124A. Here, the section base 116, the first side fairing 118A and a downstream (e.g., uncovered) portion of the trailing edge section 108 collectively form the first vane exterior surface 104A in FIGS. 3 and 4. The second side fairing 118B is laterally adjacent a second side 124B of the trailing edge section 108, and longitudinally and spanwise overlaps the trailing edge section 108 and its second side 124B. Here, the section base 116, the second side fairing 118B and the downstream portion of the trailing edge section 108 collectively form the second vane exterior surface 104B in FIGS. 3 and 4. With this arrangement, in an unlikely event of a failure event in the respective guide vane 90, air pressure in the bypass flowpath 56 will seat the trailing edge section 108 (e.g., fully) within the channel 120.
[0055] Each leading edge section 106 of FIG. 2 is moveable coupled to the trailing edge section 108, which trailing edge section 108 of FIG. 2 is a stationary section of the respective guide vane 90. Each leading edge section 106, for example, may be coupled to the trailing edge section 108 by one or more (e.g., longitudinally extending) track systems 126 and/or other guide devices. Each leading edge section 106 of FIG. 2 is operatively coupled with a vane actuator 128A, 128B (generally referred to as 128), which vane actuator 128 may be dedicated to the respective leading edge section 106 and/or independently actuatable. Each vane actuator 128 is configured to move the respective leading edge section 106 relative to the trailing edge section 108 between a (e.g., fully) retracted position (e.g., see FIG. 3) and an (e.g., fully) extended position (e.g., see FIG. 4), including one or more intermediate positions between its retracted position and its extended position. Each leading edge section 106, in particular, is configured to translate longitudinally along the respective camber line 98 between its retracted position of FIG. 3 and its extended position of FIG. 4. By moving each leading edge section 106 from (or about) the retracted position of FIG. 3 to (or towards) the extended position of FIG. 4 (or vice versa, or somewhere in between), a camber of the respective guide vane 90 along that leading edge section 106 may be changed; e.g., increased (or decreased). The vane camber describes a curvature of the respective guide vane 90. For example, as vane camber increases, a (e.g., maximum) distance 130 between the respective camber line 98 and a chord line 132 of the respective guide vane 90 may also increase.
[0056] By providing each guide vane 90 with multiple, independently adjustable leading edge sections 106, the camber of the respective guide vane 90 may be symmetrically adjusted (e.g., see FIGS. 5 and 6) or asymmetrically adjusted (e.g., see FIGS. 7 and 8) along the span line of the respective guide vane 90. For example, referring to FIG. 5, the leading edge sections 106 of a respective guide vane 90 may be disposed at their retracted positions, where respective portions of the vane leading edge 100 formed by the leading edge sections 106 are aligned. In another example, referring to FIG. 6, the leading edge sections 106 of a respective guide vane 90 may be disposed at their extended positions, where respective portions of the vane leading edge 100 formed by the leading edge sections 106 are aligned. In the arrangements of FIGS. 5 and 6, the leading edge sections 106 are aligned to provide the respective vane leading edge 100 with a substantially linear and/or continuous geometry. By contrast, referring to FIGS. 7 and 8, the leading edge sections 106 of a respective guide vane 90 may be disposed at a set of intermediate positions (e.g., partially extended/partially retracted positions) or otherwise, where respective portions of the vane leading edge 100 formed by the leading edge sections 106 are staggered. In the arrangements of FIGS. 7 and 8, the leading edge sections 106 are offset to provide the respective vane leading edge 100 with a stepped and/or non-continuous geometry. In the arrangement of FIG. 7, the inner leading edge section 106A is longitudinally recessed from the outer leading edge section 106B. Thus, the outer leading edge section 106B of FIG. 7 is moved longitudinally farther from its retracted position than the inner leading edge section 106A. In the arrangement of FIG. 8, the outer leading edge section 106B is longitudinally recessed from the inner leading edge section 106A. Thus, the inner leading edge section 106A of FIG. 8 is moved longitudinally farther from its retracted position than the outer leading edge section 106B.
[0057] The positions of the leading edge sections 106 of one, some or all of the guide vanes 90 may be adjusted individually or collectively based on and/or in response to various operational parameters and/or flight modes. Examples of the operational parameters include, but are not limited to, aircraft altitude, aircraft speed, wind speed, ambient air temperature, thrust setting, etc. Examples of the flight modes include, but are not limited to, taxing, takeoff, climb, cruise, descent, landing, etc. The positions of the leading edge sections 106 may be automatically adjusted by an onboard controller for the aircraft propulsion system 20. However, it is contemplated a pilot may also control the positions of the leading edge sections 106 in preparation or during performance of a particular maneuver. The positions of the leading edge sections 106 may be individually or collectively adjusted to improve propulsion system efficiency, shift aerodynamic loading of the propulsor rotor 42, increase or decrease conditioning of the bypass air propelled by the propulsion rotor 42, etc. Note, in some embodiments, the positions of the leading edge sections 106 may be selectively adjusted to overall shift aerodynamic loading of the propulsor rotor 42 radially inwards during one or more flight modes. An overall diameter of the propulsor rotor 42 and, thus, an overall size of the aircraft propulsion system 20 may thereby be decreased, where decreasing propulsion system size decreases overall aircraft drag and increases overall propulsion system efficiency.
[0058] Referring to FIG. 2, the leading edge sections 106 of a respective guide vane 90 may collectively form an entirety of the respective vane leading edge 100. The present disclosure, however, is not limited to such an exemplary arrangement. For example, referring to FIG. 9, the inner leading edge section 106A may be spaced radially outward from the respective vane inner end 94 and the inner case structure 52. Referring to FIG. 10, the outer leading edge section 106B may also or alternatively be spaced radially inward from the respective vane outer end 96 and the outer case structure 58. Referring to FIG. 11, the outer leading edge section 106B may also or alternatively be spaced radially outward from the inner leading edge section 106A. Moreover, while the leading edge sections 106 are symmetrically arranged along the respective vane leading edge 100 in FIG. 2, it is contemplated the leading edge sections 106 may alternatively be biased radially outward or radially inward along the span of the respective guide vane 90.
[0059] For ease of description, the guide vane structure 88 is described above as a ducted guide vane structure. However, it is contemplated the guide vane structure 88 may alternatively be configured as an open guide vane structure where, for example, the aircraft propulsion system 20 is an open rotor propulsion system and the propulsor rotor 42 is an open propulsor rotor.
[0060] While various embodiments of the present disclosure have been described, it will be apparent to those of ordinary skill in the art that many more embodiments and implementations are possible within the scope of the disclosure. For example, the present disclosure as described herein includes several aspects and embodiments that include particular features. Although these features may be described individually, it is within the scope of the present disclosure that some or all of these features may be combined with any one of the aspects and remain within the scope of the disclosure. Accordingly, the present disclosure is not to be restricted except in light of the attached claims and their equivalents.
