Outer nacelle with inlet guide vanes and acoustic treatment portion

Abstract

A gas turbine engine is provided having: a turbine; a fan rotatable by the turbine, the fan including fan blades defining an outer tip in a radial direction, a trailing edge at the tip, and a length LFB at the tip in an axial direction; and an outer nacelle surrounding the fan and at least partially surrounding the turbine, the outer nacelle including a stage of pre-swirl inlet guide vanes upstream of the fan, the stage of pre-swirl inlet guide vanes having pre-swirl inlet guide vanes, the outer nacelle further includes an inner surface in a radial direction and an acoustic treatment portion coupled to or integrated with the inner surface.

Description

Technical Field

[0001] The present disclosure relates to a gas turbine engine. Background Art

[0002] A turbofan engine typically includes a fan having a plurality of fan blades and a turbine arranged in flow communication with each other. In addition, the turbine of a turbofan engine typically includes a compressor section, a combustion section, a turbine section, and an exhaust section in a series order. In operation, air is provided from the fan to the inlet of the compressor section, where one or more axial flow compressors gradually compress the air until the compressed air reaches the combustion section. Fuel is mixed with the compressed air and combusted in the combustion section to provide combustion gases. The combustion gases are directed from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the turbine section and is then directed through the exhaust section, for example, to the atmosphere. Efficiency losses in the fan can result in a less efficient turbofan engine. BRIEF DESCRIPTION OF THE DRAWINGS

[0003] A full and enabling disclosure of the present disclosure, including the best mode thereof, to one of ordinary skill in the art is set forth in the specification with reference to the accompanying drawings, in which:

[0004] Figure 1 is a cross-sectional view of an exemplary gas turbine engine.

[0005] Figure 2 yes Figure 1 An enlarged view of the front end of an exemplary gas turbine engine showing the guide vane assembly.

[0006] Figure 3 It has multiple evenly spaced guide vane assemblies. Figure 1 An axial view of the inlet of an exemplary gas turbine engine.

[0007] Figure 4 yes Figures 1 to 3 A plan view of the inner wall of an outer nacelle of an exemplary gas turbine engine.

[0008] Figure 5 is a schematic diagram of an acoustic processing portion according to an exemplary aspect of the present disclosure.

[0009] Figure 6 is a schematic diagram of an acoustic processing portion according to another exemplary aspect of the present disclosure.

[0010] Figure 7 is a schematic diagram of an acoustic processing portion according to yet another exemplary aspect of the present disclosure.

[0011] Figure 8 is a schematic diagram of an acoustic processing portion according to yet another exemplary aspect of the present disclosure.

[0012] Figure 9 is a close-up cross-sectional view of a fan section and forward end of a turbine of a turbofan engine according to another exemplary aspect of the present disclosure.

[0013] Figure 10 is a plan view of an inner wall of an outer nacelle of a gas turbine engine according to another exemplary aspect of the present disclosure.

[0014] Figure 11 is a plan view of an inner wall of an outer nacelle of a gas turbine engine according to yet another exemplary aspect of the present disclosure.

[0015] Figure 12 is a plan view of an inner wall of an outer nacelle of a gas turbine engine according to yet another exemplary aspect of the present disclosure.

[0016] Figure 13 is a plan view of an inner wall of an outer nacelle of a gas turbine engine according to yet another exemplary aspect of the present disclosure.

[0017] Figure 14 is a plan view of an inner wall of an outer nacelle of a gas turbine engine according to yet another exemplary aspect of the present disclosure.

[0018] Figure 15 is a plan view of an inner wall of an outer nacelle of a gas turbine engine according to yet another exemplary aspect of the present disclosure.

[0019] Figure 16 is a plan view of an inner wall of an outer nacelle of a gas turbine engine according to yet another exemplary aspect of the present disclosure.

[0020] Figure 17 is a plan view of an inner wall of an outer nacelle of a gas turbine engine according to yet another exemplary aspect of the present disclosure.

[0021] Figure 18 is a plan view of an inner wall of an outer nacelle of a gas turbine engine according to yet another exemplary aspect of the present disclosure.

[0022] Figure 19 is a plan view of an inner wall of an outer nacelle of a gas turbine engine according to yet another exemplary aspect of the present disclosure.

[0023] Figure 20 is a plan view of an inner wall of an outer nacelle of a gas turbine engine according to yet another exemplary aspect of the present disclosure.

[0024] Figure 21 is a plan view of an inner wall of an outer nacelle of a gas turbine engine according to yet another exemplary aspect of the present disclosure.

[0025] Figure 22 is a plan view of an inner wall of an outer nacelle of a gas turbine engine according to yet another exemplary aspect of the present disclosure.

[0026] Figure 23 is a plan view of an inner wall of an outer nacelle of a gas turbine engine according to yet another exemplary aspect of the present disclosure. DETAILED DESCRIPTION

[0027] Reference will now be made in detail to the present embodiments of the present disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter references to refer to features in the drawings. Like or similar reference numerals have been used in the drawings and the description to refer to like or similar parts of the present disclosure.

[0028] The word "exemplary" is used herein to mean "serving as an example, instance, or illustration." Any embodiment described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other embodiments. Additionally, unless expressly stated otherwise, all embodiments described herein should be considered exemplary.

[0029] For the purposes of this description, the terms "upper," "lower," "right," "left," "vertical," "horizontal," "top," "bottom," "lateral," "longitudinal" and their derivatives will refer to the disclosure as oriented in the accompanying drawings.

[0030] As may be used herein, the terms “first,” “second,” “third,” and other ordinal numbers are used to distinguish one component from another and are not intended to indicate the position or importance of each component.

[0031] The terms "fore" and "aft" refer to relative positions within a gas turbine engine, with "fore" referring to a position closer to the engine inlet and "aft" referring to a position closer to the engine nozzle or exhaust.

[0032] The terms "upstream" and "downstream" refer to relative directions relative to the flow of fluid in a fluid path. For example, "upstream" refers to the direction from which the fluid is flowing, and "downstream" refers to the direction toward which the fluid is flowing.

[0033] The term "attached" refers to two components that are directly connected to each other. The term "integrated" refers to two components that are formed simultaneously as a single piece, or two components that are formed separately and subsequently secured to each other. The term "unitary structure" refers to a one-piece structure that is formed integrally, such that the components of the unitary structure are formed simultaneously.

[0034] The singular forms "a," "an," and "the" include plural referents unless the context clearly dictates otherwise.

[0035] In a context such as “at least one of A, B, and C,” the term “at least one of” means only A, only B, only C, or any combination of A, B, and C.

[0036] The phrases "from X to Y" and "between X and Y" each refer to a range of values ​​including the endpoints (ie, to a range that includes both X and Y values).

[0037] Typically, a gas turbine engine includes a fan and a turbine, where the turbine rotates the fan to generate thrust. The turbine includes a compressor section, a combustion section, a turbine section, and an exhaust section, and defines a working gas flow path therethrough. For ducted gas turbine engines, the gas turbine engine also includes an outer nacelle surrounding at least a portion of the fan and turbine.

[0038] Typically, the efficiency of a gas turbine engine can be improved by increasing the size of the fan. However, for larger fans, particularly for larger fans in direct-drive gas turbine engines (gas turbine engines without a reduction gearbox mechanically located between the drive turbine and the fan), the tips of the fan blades of the fan can rotate at a relatively high speed. In this manner, the fan can be referred to as a "high tip speed fan." In order to minimize airflow separation or other negative aerodynamic effects at the outer ends of the fan blades of a high tip speed fan, the inventors have discovered that pre-swirl inlet guide vanes (IGVs) can be provided with a nacelle at a position upstream of the fan.

[0039] The inventors of the present disclosure sought a device incorporating a pre-swirl IGV that would result in a net gain in achieving the aforementioned goals.

[0040] In particular, the present inventors have recognized that including a pre-swirl IGV with an outer nacelle would allow a direct drive gas turbine engine to increase the diameter of the fan while avoiding or reducing negative aerodynamic effects at the outer ends of the fan blades of the fan.

[0041] However, the present inventors have discovered that combining pre-swirl IGVs with larger fans creates an opportunity for increased noise due to airflow pressure fluctuations associated with the fan blade pass frequency of the fan. The present inventors have unexpectedly discovered that, in some arrangements, including an acoustic treatment to the outer nacelle can provide the benefits associated with including pre-swirl IGVs in direct-drive gas turbine engines with larger fans / high tip speed fans without an excessive increase in noise.

[0042] In particular, the present inventors unexpectedly discovered that, in designing gas turbine engines having high tip speed fans, outer nacelles with pre-swirl IGVs to accommodate the high tip speed fans, and increased acoustic treatments to address noise generation, contrary to previous belief and expectation, the costs associated with including high tip speed fans and outer nacelles with pre-swirl IGVs (e.g., noise) were overcome in at least some designs by gas turbine engine efficiency benefits. In particular, during the design of several gas turbine engines having high tip speed fans and outer nacelles with pre-swirl IGVs and acoustic treatments of varying thrust ratings and varying fan sizes / fan tip speeds (including the configurations shown and described in detail herein), the present inventors discovered a relationship between fan blade and pre-swirl inlet guide vane spacing and acoustic treatment length such that including a high tip speed fan and an outer nacelle with pre-swirl IGVs and acoustic treatments according to one or more exemplary aspects described herein resulted in a net benefit to the overall engine design.

[0043] In order to achieve the goal of an improved gas turbine engine capable of providing the desired aerodynamic efficiency gains associated with operating a fan at higher speeds, the inventors proceeded in the process of designing several different types of outer nacelles having a pre-swirl IGV and an acoustic treatment portion in combination therewith (including the outer nacelle described herein, which will be described in more detail below) a gas turbine engine having an outer nacelle having a pre-swirl IGV and an acoustic treatment portion in combination therewith, having various fan blade and pre-swirl inlet guide vane spacings and acoustic treatment lengths; examining the operability and aerodynamic efficiency characteristics of the designed gas turbine engine; redesigning the gas turbine engine to change the parameters based on the impact on other aspects of the gas turbine engine; re-examining the operability and aerodynamic efficiency characteristics of the redesigned gas turbine engine; and so on.

[0044] Referring now to the drawings, in which like numerals represent like elements throughout, Figure 1 is a cross-sectional view of a gas turbine engine according to an exemplary embodiment of the present disclosure. More specifically, Figure 1 In an embodiment of the present invention, the gas turbine engine is an aviation turbofan jet engine, referred to herein as a "turbofan engine 10." The turbofan engine 10 is configured to be mounted to an aircraft, such as in an underwing configuration or a tail-mounted configuration. Figure 1As shown, the turbofan engine 10 defines an axial direction A (extending parallel to a longitudinal centerline provided for reference), a radial direction R, and a circumferential direction (e.g., a direction extending about the axial direction A). The longitudinal centerline defines a longitudinal centerline 12 of the turbofan engine 10. Generally, the turbofan engine 10 includes a fan section 14 and a turbine 16 disposed downstream of the fan section 14 (the turbine 16 is sometimes also referred to as a "core turbine engine," or alternatively, a "core turbine engine").

[0045] The exemplary turbine 16 shown generally includes a substantially tubular outer casing 18 defining an annular inlet 20. Outer casing 18 encloses, in series flow relationship, a compressor section comprising a first supercharger or low-pressure (LP) compressor 22 and a second high-pressure (HP) compressor 24; a combustion section 26; a turbine section comprising a first high-pressure (HP) turbine 28 and a second low-pressure (LP) turbine 30; and an exhaust nozzle section 32. A high-pressure (HP) shaft drivingly connects the HP turbine 28 to the HP compressor 24. A low-pressure (LP) shaft 36 drivingly connects the LP turbine 30 to the LP compressor 22. The compressor section, combustion section 26, turbine section, and exhaust nozzle section 32 are arranged in series flow order and, together, define a core air flow path 37 through the turbine 16. It is also contemplated that the present disclosure is compatible with engines having intermediate-pressure turbines, for example, engines having three spools.

[0046] Still refer to Figure 1 In the exemplary embodiment, fan section 14 includes a single-stage variable-pitch fan 38, to which turbine 16 is operably coupled for driving fan 38. Fan 38 includes a plurality of rotatable fan blades 40 coupled to a disk 42 in a spaced-apart manner. As shown, fan blades 40 generally extend outwardly from disk 42 in a radial direction R. Each fan blade 40 is rotatable relative to disk 42 about a pitch axis P by virtue of fan blades 40 being operably coupled to a suitable actuating member 44. Actuating member 44 is configured to collectively change the pitch of fan blades 40, for example, to change the pitch of fan blades 40 in unison. Fan blades 40, disk 42, and actuating member 44 are rotatable together about longitudinal centerline 12 via LP shaft 36. Notably, in the illustrated embodiment, fan 38 and fan blades 40 can be rotated by LP shaft 36 via a 1:1 mechanical connection (i.e., without a reduction gearbox or other speed change mechanism). Therefore, it should be understood that the exemplary turbofan engine 10 illustrated is configured to directly drive a gas turbine engine.

[0047] In an exemplary embodiment, fan section 14 includes twenty-two (22) or fewer fan blades 40. In other exemplary embodiments, fan section 14 includes a different number of fan blades 40, such as twenty (20), eighteen (18), sixteen (16), or another number of fan blades 40.

[0048] Still refer to Figure 1 In the exemplary embodiment, disk 42 is covered by a rotatable forward nacelle or hub 48 having an aerodynamic profile to facilitate airflow over the plurality of fan blades 40. Additionally, exemplary fan section 14 includes an annular fan case or outer nacelle 50 that at least partially (and, for the illustrated embodiment, circumferentially) surrounds fan 38 and at least a portion of turbine 16.

[0049] More specifically, the outer nacelle 50 includes an inner wall 52 (defining an inner surface; not separately labeled), and a downstream section 54 of the inner wall 52 of the outer nacelle 50 extends above an outer portion of the turbine 16 to define a bypass airflow passage 56 therebetween. Additionally, for the illustrated embodiment, the outer nacelle 50 is supported relative to the turbine 16 by a plurality of circumferentially spaced outlet guide vanes 55. The outer nacelle 50 includes an inlet 60 at a leading edge 61 of the outer nacelle 50.

[0050] During operation of turbofan engine 10, a volume of air 58 enters turbofan engine 10 through outer nacelle 50 and / or inlet 60 of fan section 14. As volume of air 58 passes through fan blades 40, a first portion of air 58, as indicated by arrow 62, is directed or channeled into bypass airflow passage 56, and a second portion of air 58, as indicated by arrow 64, is directed or channeled into core air flow path 37. The pressure of the second portion of air, indicated by arrow 64, then increases as it is directed through HP compressor 24 and into combustion section 26, where it is mixed with fuel and combusted to provide combustion gases 66. Combustion gases 66 are directed from combustion section 26 through HP turbine 28. In HP turbine 28, a portion of the thermal and / or kinetic energy from combustion gases 66 is extracted via successive stages of HP turbine stator blades 68 coupled to outer casing 18 and HP turbine rotor blades 70 coupled to high pressure (HP) shaft 34, thereby causing HP shaft 34 to rotate, thereby supporting operation of HP compressor 24. Combustion gases 66 are then directed through LP turbine 30, wherein a second portion of the thermal and / or kinetic energy is extracted from combustion gases 66 via successive stages of LP turbine stator blades 72 coupled to outer casing 18 and LP turbine rotor blades 74 coupled to LP shaft 36, thereby causing LP shaft 36 to rotate, thereby supporting operation of LP compressor 22 and / or rotation of fan 38.

[0051] The combustion gases 66 are then directed through the exhaust nozzle section 32 of the turbine 16 to provide propulsive thrust. At the same time, the pressure of the first portion of air 62 is significantly increased as the first portion of air 62 is directed through the bypass airflow passage 56 before being discharged from the fan nozzle exhaust section 76 of the turbofan engine 10, also providing propulsive thrust. The HP turbine 28, the LP turbine 30, and the exhaust nozzle section 32 at least partially define a hot gas path 78 for directing the combustion gases 66 through the turbine 16.

[0052] In some exemplary embodiments, the exemplary turbofan engine 10 of the present disclosure may be a relatively high-power class turbofan engine 10. Thus, when operating at rated speed, the turbofan engine 10 may be configured to generate a relatively large amount of thrust. More specifically, when operating at rated speed, the turbofan engine 10 may be configured to generate at least 20,000 pounds of thrust, such as at least approximately 25,000, 30,000, and up to, for example, 150,000 pounds of thrust. Thus, the turbofan engine 10 may be referred to as a relatively high-power class gas turbine engine.

[0053] Still refer to Figure 1 , and as mentioned before, Figure 1 The exemplary turbofan engine 10 shown in FIG is configured as a direct drive turbofan engine 10 , however, other configurations are also contemplated. For example, a geared configuration of the turbofan engine 10 may be implemented, wherein a reduction gear ratio may be selected to maintain high speed operation of the fan 38 .

[0054] However, still with reference to the direct drive turbofan engine 10, the interaction between the fan 38 and the power generation components in the turbine 16, where the relative speeds are fixed relative to each other, should be considered. In order to improve the efficiency of the turbine 16, the LP turbine 30 is configured to rotate at a relatively high rotational speed. In the case of the direct drive configuration, this relatively high speed of rotation of the turbine 16 also causes the plurality of fan blades 40 of the fan 38 to rotate at a relatively high rotational speed. For example, during operation of the turbofan engine 10 at rated speed, the fan tip speed of each of the plurality of fan blades 40 is greater than 1,250 feet per second. In certain exemplary embodiments, during operation of the turbofan engine 10 at rated speed, the fan tip speed of each of the plurality of fan blades 40 may be greater than 1,350 feet per second, such as greater than 1,450 feet per second, such as greater than 1,550 feet per second, such as up to 2,200 feet per second.

[0055] Despite these relatively high fan tip speeds, fan 38 is still designed to define a relatively low fan pressure ratio. For example, during operation of turbofan engine 10 at rated speed, the fan pressure ratio of fan 38 is greater than 1.0 and less than 1.5. For example, during operation of turbofan engine 10 at rated speed, the fan pressure ratio may be between approximately 1.15 and approximately 1.5, such as between approximately 1.25 and approximately 1.4.

[0056] As will be appreciated, operating the high-speed turbofan engine 10 in this manner typically results in a loss of efficiency of the fan 38 due to shock losses and flow separation, particularly at the outer tips of the plurality of fan blades 40 of the fan 38 along the radial direction R. Therefore, as will be described in greater detail below, the turbofan engine 10 also includes one or more inlet pre-swirl features 80 located upstream of the plurality of fan blades 40 of the fan 38 to offset or minimize such efficiency losses of the fan 38. By including such inlet pre-swirl features, the efficiency gains of the turbine 16 due to, for example, the increased rotational speed of the LP turbine 30 may outweigh the potential efficiency losses identified above.

[0057] Now also refer to Figure 2 , provides Figure 1 FIG2 is a close-up cross-sectional view of the forward end of the fan section 14 and turbine 16 of an exemplary turbofan engine 10. As described above, the turbofan engine 10 includes pre-swirl inlet guide vanes 100, and more specifically, a stage of pre-swirl inlet guide vanes 100. The stage of pre-swirl inlet guide vanes 100 is located upstream of the plurality of fan blades 40 of the fan 38 and is attached to or integrated into the body 51 of the nacelle 50.

[0058] More specifically, for Figure 1 and Figure 2 In an embodiment of the present invention, each of the plurality of pre-swirl inlet guide vanes is a partial span inlet guide vane cantilevered from the body 51 of the nacelle 50 at a position forward of the plurality of fan blades 40 of the fan 38 along the axial direction A. In this configuration, each of the plurality of pre-swirl inlet guide vanes 100 defines a base 102 along the radial direction R (see FIG. Figure 2 ), and are attached or connected to the body 51 of the nacelle 50 at the base 102. For example, each of the inlet pre-swirl features 80 may be bolted to the body 51 of the nacelle 50 at the base 102, welded to the body 51 of the nacelle 50 at the base 102, or attached to the body 51 of the nacelle 50 at the base 102 in any other suitable manner.

[0059] Furthermore, for the embodiment shown, each of the plurality of pre-swirl inlet guide vanes 100 extends generally along the radial direction R from its respective base 102 to a respective tip 104 generally along the radial direction R. Furthermore, as will be appreciated, for the embodiment shown, each of the plurality of pre-swirl inlet guide vanes 100 is disconnected from an adjacent one of the plurality of pre-swirl inlet guide vanes 100 at its respective tip 104. More specifically, for the embodiment shown, each pre-swirl inlet guide vane 100 is fully supported by its connection to, or integration with, the main body 51 of the nacelle 50 at its respective base 102 (rather than through any structure extending between adjacent inlet pre-swirl features at a location inboard of the outer end along the radial direction R).

[0060] like Figure 2 As shown, each of the plurality of pre-swirl inlet guide vanes 100 does not fully extend between the nacelle 50 and, for example, the hub 48. More specifically, for the illustrated embodiment, each of the plurality of pre-swirl inlet guide vanes 100 defines a pre-swirl feature span 106 along the radial direction R. More specifically, each of the plurality of pre-swirl inlet guide vanes 100 further defines a leading edge 108 and a trailing edge 110, wherein the pre-swirl feature span 106 refers to a measurement along the radial direction R between the base 102 and the tip 104 of the pre-swirl feature 80 at the leading edge 108 of the pre-swirl feature 80. Similarly, it should be understood that each of the plurality of fan blades 40 of the fan 38 defines a fan blade span 112 along the radial direction R. More specifically, each of the plurality of fan blades 40 of fan 38 defines a leading edge 114 and a trailing edge 116 , where fan blade span 112 refers to a measurement along radial direction R between the radially outer tip 82 and base 84 of the fan blade 40 at the leading edge 114 of the respective fan blade 40 .

[0061] For the illustrated embodiment, the pre-swirl signature span 106 is at least approximately five percent of the fan blade span 112 and up to approximately fifty-five percent of the fan blade span 112. For example, in certain exemplary embodiments, the pre-swirl signature span 106 may be between approximately fifteen percent of the fan blade span 112 and approximately forty-five percent of the fan blade span 112, such as between approximately thirty percent of the fan blade span 112 and approximately forty percent of the fan blade span 112.

[0062] Although not shown, in certain exemplary embodiments, the number of the plurality of pre-swirl inlet guide vanes 100 may be substantially equal to the number of fan blades 40 of the fan 38 of the turbofan engine 10. However, in other embodiments, the number of the plurality of pre-swirl inlet guide vanes 100 may be greater than the number of fan blades 40 of the fan 38 of the turbofan engine 10, or alternatively, may be less than the number of fan blades 40 of the fan 38 of the turbofan engine 10.

[0063] Furthermore, it should be understood that in other exemplary embodiments, the turbofan engine 10 may include any other suitable number of inlet pre-swirl features 80 and / or circumferential spacing of inlet pre-swirl features 80. For example, the turbofan engine 10 may include fewer than fifty and at least eight inlet pre-swirl features 80.

[0064] In addition, a brief reference Figure 3 , provides Figure 1 and Figure 2 The axial view of the inlet 60 of the turbofan engine 10 includes a plurality of pre-swirl inlet guide vanes 100 arranged circumferentially around the nacelle 50. More specifically, Figure 3 The number of pre-swirl inlet guide vanes 100 shown in FIG ranges from about ten pre-swirl inlet guide vanes 100 to about fifty pre-swirl inlet guide vanes 100. The plurality of pre-swirl inlet guide vanes 100 are substantially evenly arranged along the circumferential direction C. More specifically, each of the plurality of pre-swirl inlet guide vanes 100 defines a circumferential spacing 127 with an adjacent pre-swirl inlet guide vane 100, and each of the circumferential spacings 127 is substantially equal to one another.

[0065] However, it should be understood that in other exemplary embodiments, the plurality of pre-swirl inlet guide vanes 100 may be unevenly spaced along the circumferential direction C.

[0066] Still refer to Figure 2 In the embodiment of the present invention, it will be appreciated that each of the plurality of pre-swirl inlet guide vanes 100 is configured to pre-swirl air 58 provided through the inlet 60 of the nacelle 50 upstream of the plurality of fan blades 40 of the fan 38. As described above, pre-swirl the air 58 provided through the inlet 60 of the nacelle 50 before such air 58 reaches the plurality of fan blades 40 of the fan 38 may reduce separation losses and / or impulse losses, allowing the fan 38 to operate at the aforementioned relatively high fan tip speed with minimal loss of efficiency.

[0067] The inventors have discovered several aspects of the pre-swirl inlet guide vanes 100 and fan blades 40 that affect the operation and efficiency of the turbofan engine 10. For example, Figure 2As shown, fan blade 40 includes a leading edge 114 and a trailing edge 116, and a tip 82 along a radial direction R. Fan blade 40 also defines a length L along an axial direction A from leading edge 114 to trailing edge 116 at tip 82. FB .

[0068] Similarly, the illustrated pre-swirl inlet guide vane 100 among the plurality of pre-swirl inlet guide vanes 100 defines a 15% span position 118 (where 0% span is at the base 102 and 100% span is at the tip 104). In other words, the 15% span position 118 represents a position that is 15% of the distance traveled from the base 102 to the tip 104. The pre-swirl inlet guide vane 100 defines a length L along the axial direction A at the 15% span position 118. IGV .

[0069] Furthermore, the turbofan engine 10 defines a length L along the axial direction A from the leading edge 108 of the pre-swirl inlet guide vane 100 at the 15% span position 118 to the trailing edge 116 of the fan blade 40 at the tip of the fan blade 40 S .

[0070] Furthermore, as will be appreciated from the description herein, while including a plurality of pre-swirl inlet guide vanes 100 improves the efficiency of the exemplary turbofan engine 10 illustrated, positioning the plurality of pre-swirl inlet guide vanes 100 upstream of the fan blades 40 of the fan 38 creates an opportunity for increased noise due to airflow pressure fluctuations associated with the passing frequency of the fan blades 40 of the fan 38. To avoid excessive noise generation resulting from including the pre-swirl inlet guide vanes 100, the exemplary turbofan engine 10 illustrated, and more specifically, the outer nacelle 50 of the illustrated turbofan engine 10, further includes an acoustic treatment 120. The present inventors have unexpectedly discovered that, in some arrangements, including an acoustic treatment 120 on the outer nacelle 50 can provide the benefits associated with including the pre-swirl inlet guide vanes 100 in the turbofan engine 10 of the present disclosure without an excessive increase in noise.

[0071] As used herein, the term "acoustic treatment" refers to any material or system applied to or integrated into the inner wall 52 of the outer nacelle 50 that is designed or selected for its properties to absorb, dissipate, or reduce the transmission of sound energy. The term acoustic treatment includes materials, structures, composite materials, and combinations thereof that can attenuate airborne noise, structure-borne vibrations, or any combination thereof, thereby mitigating one or more aspects of noise emissions from the engine. The term includes, but is not limited to, porous materials, foam-based materials, damping materials, and structures or devices that exhibit sound absorbing properties or modify the acoustic environment to effectively reduce the perception of noise (including those described below with reference to Figures 5 to 8 structure described).

[0072] In particular, now also refer to Figure 4 , provides Figures 1 to 3 FIG. 1 is a plan view of the inner wall 52 of the outer nacelle 50, wherein the stages of pre-swirl inlet guide vanes 100 are arranged along the circumferential direction C, showing the acoustic treatment portion 120 of the outer nacelle 50. In the illustrated embodiment, the acoustic treatment portion 120 extends from a forward end 132 to an aft end 134, wherein the forward end 132 is positioned forward of the stages of pre-swirl inlet guide vanes 100, and the aft end 134 is positioned behind the plurality of fan blades 40 of the fan (at the rear of the fan). Figure 4 The acoustic treatment portion 120 defines a length L along the axial direction A from the front end 132 to the rear end 134. AT .

[0073] In this manner, for the exemplary embodiment shown, the acoustic treatment 120 is positioned at least along the axial direction A between the stages of the pre-swirl inlet guide vanes 100 and the tip 82 of the fan blade 40, and is also positioned partially along the axial direction A at the stages of the pre-swirl inlet guide vanes 100, and at least partially along the axial direction A at the tip 82 of the fan blade 40. It should be understood that, as used in this context, the term "at" means that a first component (e.g., the acoustic treatment 120) has at least a portion that overlaps along the axial direction A with at least a portion of a second component (e.g., the pre-swirl inlet guide vanes 100 or the tip 82 of the fan blade 40).

[0074] In addition, still refer to Figure 4 It should be understood that the exemplary acoustic treatment 120 extends continuously along the circumference of the inner wall 52 of the outer nacelle 50 (eg, extends continuously in the circumferential direction C).

[0075] In at least some exemplary embodiments, the acoustic treatment 120 may include a perforated plate 122 having a hollow body 124. Specifically, referring now to Figure 5 , provides a schematic cross-sectional view of the acoustic treatment portion 120 according to an exemplary aspect of the present disclosure. Figure 5 As shown in the embodiment of FIG. 1 , the exemplary acoustic treatment portion 120 includes a perforated plate 122 and a hollow body 124. The hollow body 124 includes a liner 128 that defines an interior void adjacent to the perforated plate 122. The perforated plate 122 defines a plurality of openings 126, thereby allowing the external environment to communicate with the interior void of the hollow body 124. Sound waves can enter the hollow body 124 through the plurality of openings 126, allowing for attenuation of noise generated by the orientation of the exit guide vanes 55 in the forward swept arrangement.

[0076] The perforated plate 122 is coupled to the liner 128 by a plurality of extensions 130 extending from the perforated plate 122 to the liner 128. In certain exemplary embodiments, the acoustic treatment 120 may also include additional structure to increase the noise attenuation achieved by the acoustic treatment 120 at desired frequencies. The additional structure may be a wall or other extension 130 (shown in phantom) extending from the perforated plate 122, from the liner 128, or both; may be perforations in the wall or extension 130; may be additional or alternative wall or extension 130; and so forth.

[0077] Now refer to Figure 6 , provides a schematic top view of a perforated plate 122 according to an exemplary aspect of the present disclosure. The perforated plate 122 can include a plurality of openings 126 spaced apart in a uniform manner.

[0078] Now refer to Figure 7 , provides a schematic top view of a perforated plate 122 according to another exemplary aspect of the present disclosure. Figure 7 It will be appreciated from the view of FIG. 1 that the plurality of openings 126 of the perforated plate 122 may define a non-circular shape, such as an elongated or oval shape.

[0079] Now refer to Figure 8 , provides a schematic top view of a perforated plate 122 according to yet another exemplary embodiment of the present disclosure. Figure 8 As can be appreciated from the view of FIG, the plurality of openings 126 can define non-uniform sizes and non-uniform spacings. Such a configuration can, for example, allow the acoustic treatment 120 to target noise of various frequencies.

[0080] As previously mentioned, the present inventors unexpectedly discovered that, in the process of designing a gas turbine engine having a high tip speed fan and an outer nacelle with a pre-swirl IGV and an acoustic treatment—that is, designing a gas turbine engine having a high tip speed fan and an outer nacelle with a pre-swirl IGV and an acoustic treatment and evaluating overall engine and aerodynamic efficiency performance—a significant relationship between the fan blade and pre-swirl inlet guide vane spacing and the acoustic treatment length was found. This relationship can be considered an indicator of the ability of a gas turbine having a high tip speed fan and an outer nacelle with a pre-swirl IGV and an acoustic treatment to achieve improvements in aerodynamic performance at the outer tips of the fan blades during high-speed operation without incurring an excessive increase in noise generation.

[0081] This relationship applies to gas turbine engines having an outer nacelle with a pre-swirl IGV and, for example, an acoustic treatment integrated with the inner surface of the outer nacelle. This relationship connects the fan blade tip length, the pre-swirl inlet guide vane length, the fan blade and pre-swirl inlet guide vane spacing, and the acoustic treatment length, as described in more detail below.

[0082] In particular, the inventors have discovered that the inclusion of a pre-swirl IGV upstream of a high tip speed fan can mitigate aerodynamic issues at the outer end of the fan as the pre-swirl IGV interacts with fluctuations in pressure from the fan blades as a result of the fan blades' rotation and passing over the fan blades at a blade-by frequency. The inventors have discovered that the amplitude of the pressure fluctuations is related to the axial length of the tips of the fan blades (fan blade tip length), with higher fan blade tip lengths resulting in higher amplitude pressure fluctuations.

[0083] Furthermore, the present inventors have found that the axial length of the pre-swirl inlet guide vanes, ie the pre-swirl inlet guide vane length, affects the level of interaction with pressure fluctuations, wherein a higher pre-swirl inlet guide vane length results in greater interaction (and more noise generation).

[0084] Similarly, the present inventors have discovered that the axial spacing between the fan blade trailing edge and the pre-swirl inlet guide vane leading edge, i.e., the fan blade to pre-swirl inlet guide vane spacing, also affects the level of interaction with pressure fluctuations, with lower fan blade to pre-swirl inlet guide vane spacing resulting in greater interaction (and more noise generation).

[0085] Furthermore, the inventors have discovered that the amount of acoustic treatment applied to the outer nacelle, and more specifically, its length, influences the level of noise attenuation achieved by the treatment, with longer treatments resulting in greater noise attenuation. However, as the length of the treatment increases, the cost of manufacturing the part increases, and further, the sound attenuation benefits have diminishing returns.

[0086] Thus, the relationships discovered below can identify a gas turbine engine with a high tip speed fan and an outer nacelle with pre-swirl IGVs and acoustic treatment that is capable of achieving the desired aerodynamic efficiency gains associated with operating the fan at higher speeds while avoiding excessive increases in noise generation, and that is suitable for specific mission requirements that take into account efficiency, weight, noise, complexity, reliability, and other factors that influence the optimal selection of a gas turbine engine with a high tip speed fan and an outer nacelle with pre-swirl IGVs and acoustic treatment.

[0087] In addition to producing an improved gas turbine engine having the aforementioned features, as explained in detail above, by utilizing this relationship, the inventors have discovered that the number of suitable or feasible gas turbine engine designs that can meet the aforementioned design requirements can be significantly reduced, thereby facilitating more rapid selection of a design for consideration as a gas turbine engine under development (having a high tip speed fan and an outer nacelle with pre-swirl IGVs and acoustic treatments). This benefit provides greater understanding of the requirements for a given gas turbine engine before the specific technical, integration, and system requirements are fully developed. This benefit avoids late-stage redesigns.

[0088] The present inventors have found that the desired relationship provided for the improved gas turbine engine provides a length L in the axial direction of the acoustic treatment of AT In particular, the relationship is as follows:

[0089] (L IGV 2 / L S )×UCF1 <L AT <(L FB 3 / L IGV )×UCF2 (1)

[0090] The length L IGV The length L is the axial length of the pre-swirl inlet guide vane at the 15% span position, in inches; S is the length in inches along the axial direction from the leading edge of the pre-swirl inlet guide vane at the 15% span position to the trailing edge of the fan blade at the tip of the fan blade; length L FB is the length of the fan blade in the axial direction at the tip of the fan blade divided by the length in inches; UCF1 is equal to 1 -1 The first unit correction factor of inches; and UCF2 is equal to 1 -2 Second unit correction factor for inches.

[0091] In this way, it can be understood that the above relationship (1) requires the length L of the acoustic treatment portion to be AT Greater than length L IGV The square of the length L S Multiply by the first unit correction factor and less than the length L FB cubed divided by the length L IGV Multiply by the second unity correction factor.

[0092] In particular, the above-described benefits apply to gas turbine engines having the values ​​summarized in Table 1 below.

[0093]

[0094] In addition, the following are one or more example gas turbine engines of the present disclosure having an outer nacelle with a pre-swirl IGV and an acoustic treatment portion. In particular, each of the gas turbine engines defines a length L in the axial direction of the pre-swirl inlet guide vane at a 15% span position. IGV , in inches; the length L along the axial direction from the leading edge of the pre-swirl inlet guide vane at the 15% span position to the trailing edge of the fan blade at the tip of the fan blade S , in inches; the length L along the axial direction of the fan blade at the tip of the fan blade divided by the length FB , in inches; and the length L along the axial direction of the acoustic treatment portion AT .

[0095] Example 1: In a first example gas turbine engine having an outer nacelle with a pre-swirl IGV and acoustic treatment, the gas turbine engine defines a length L equal to 7.2 inches (in) FB ; Length L equal to 0.4in IGV ; Length L equal to 7.7in S ; and a length L greater than 0.02 in and less than 1037 in AT .

[0096] Example 2: In a first example gas turbine engine having an outer nacelle with a pre-swirl IGV and an acoustic treatment, the gas turbine engine defines a length L equal to 7.2 inches (in) FB ; Length L equal to 2.4in IGV ; Length L equal to 10.8in S ; and a length L greater than 0.52 in and less than 157.5 in AT .

[0097] Example 3: In a first example gas turbine engine having an outer nacelle with a pre-swirl IGV and an acoustic treatment, the gas turbine engine defines a length L equal to 7.2 inches (in) FB ; Length L equal to 3.6in IGV ; Length L equal to 12.6in S ; and a length L greater than 1.0 in and less than 103.7 in AT .

[0098] Example 4: In a first example gas turbine engine having an outer nacelle with a pre-swirl IGV and an acoustic treatment, the gas turbine engine defines a length L equal to 7.2 inches (in) FB ; Length L equal to 5.4in IGV; Length L equal to 15.3in S ; and a length L greater than 1.9 in and less than 69.1 in AT .

[0099] Example 5: In a first example gas turbine engine having an outer nacelle with a pre-swirl IGV and an acoustic treatment, the gas turbine engine defines a length L equal to 7.2 inches (in) FB ; Length L equal to 7.2in IGV ; Length L equal to 18.0 in S ; Length L greater than 2.9in and less than 51.8in AT .

[0100] Example 6: In a first example gas turbine engine having an outer nacelle with a pre-swirl IGV and an acoustic treatment, the gas turbine engine defines a length L equal to 7.2 inches (in) FB ; Length L equal to 9 inches IGV ; Length L equal to 20.7in S ; Length L greater than 3.9in and less than 41.5in AT .

[0101] However, it should be understood that the above reference Figures 1 to 8 The described exemplary embodiments are provided by way of example only. In other exemplary embodiments, turbofan engine 10 , including outer nacelle 50 and acoustic treatment 120 of outer nacelle 50 , may be arranged in any other suitable manner.

[0102] For example, now refer to Figure 9 , provides a close-up cross-sectional view of the forward end of the fan section 14 and turbine 16 of a turbofan engine 10 according to another exemplary aspect of the present disclosure. Figure 9 The exemplary turbofan engine 10 of can be configured in a similar manner to the exemplary turbofan engine 10 described above (see, e.g., Figure 2 ).

[0103] Thus, it should be understood that the exemplary turbofan engine 10 generally includes a fan section 14 having a fan 38, a turbine 16, and an outer nacelle 50 surrounding at least a portion of the fan 38 and the turbine 16. The outer nacelle 50 includes a stage of pre-swirl inlet guide vanes 100 in front of the fan 38 and an acoustic treatment 120. However, for Figure 9 In the embodiment of the present invention, the acoustic treatment portion 120 further extends to the pre-swirl inlet guide vanes 100 of the stage of the swirl inlet guide vanes 100. In particular, the acoustic treatment portion 120 includes an inlet guide vane section 121 coupled to or integrated with the pre-swirl inlet guide vanes 100.

[0104] In the illustrated embodiment, the inlet guide vane section 121 extends along the length of the pre-swirl characteristic span 106 of the pre-swirl inlet guide vane 100 (e.g., at least 90% of the pre-swirl characteristic span 106 of the pre-swirl inlet guide vane 100 ) and also extends along the length of the chord 135 of the pre-swirl inlet guide vane 100 (e.g., at least 90% of the chord 135 of the pre-swirl inlet guide vane 100 ).

[0105] In some embodiments, the inlet guide vane section 121 may extend completely around the pre-swirl inlet guide vane 100 , for example, on both the pressure and suction sides of the pre-swirl inlet guide vane 100 .

[0106] Furthermore, as indicated by the dashed lines, in certain exemplary embodiments, the inlet guide vane section 121 may extend only partially along the pre-swirl characteristic span 106 of the pre-swirl inlet guide vane 100 (e.g., at least 10%, and up to 75%, such as up to 50% of the pre-swirl characteristic span 106 of the pre-swirl inlet guide vane 100; ending at reference line 137, provided in dashed lines).

[0107] Furthermore, as also shown by the dashed lines, in certain exemplary embodiments, the inlet guide vane section 121 of the acoustic treatment portion 120 may extend only partially along the chord 135 of the pre-swirl inlet guide vane 100 and may be located closer to the leading edge 108 of the pre-swirl inlet guide vane 100, or alternatively, closer to the trailing edge 110 of the pre-swirl inlet guide vane 100 (e.g., starting or ending at the reference line 139 shown by the dashed lines).

[0108] In certain exemplary embodiments, inlet guide vane section 121 may be located in any one or more quadrants formed by reference lines 137 , 139 .

[0109] Now generally refer to Figures 10 to 23 , various other exemplary embodiments of the present disclosure are provided. Figures 10 to 23 Each provides a plan view of the inner wall 52 of the outer nacelle 50 of the turbofan engine 10 according to an exemplary aspect of the present disclosure, depicting a stage of pre-swirl inlet guide vanes 100 arranged along the circumferential direction C and an acoustic treatment 120 of the outer nacelle 50 . Figures 10 to 23 Each of the above also depicts the length L of the acoustic treatment 120 along the axial direction A of the corresponding turbofan engine 10. AT , the length L AT Defined between a front end 132 of the acoustic treatment 120 and a rear end 134 of the acoustic treatment 120 .

[0110] Special References Figure 10The acoustic treatment portion 120 is located behind the stage of pre-swirl inlet guide vanes 100 , wherein the front end 132 of the acoustic treatment portion 120 is located behind the trailing edge 110 of each pre-swirl inlet guide vane 100 of the stage of pre-swirl inlet guide vanes 100 .

[0111] Instead, now with special reference Figure 11 The acoustic treatment portion 120 is located in front of the stage of pre-swirl inlet guide vanes 100 , wherein the rear end 134 of the acoustic treatment portion 120 is located in front of the leading edge 108 of each pre-swirl inlet guide vane 100 of the stage of pre-swirl inlet guide vanes 100 .

[0112] In addition, special reference Figure 12 The acoustic treatment portion 120 includes a plurality of acoustic treatment panels 136 spaced apart along the axial direction A. More specifically, for the embodiment shown, the plurality of acoustic treatment panels 136 include a first front acoustic treatment panel 136A and a second rear acoustic treatment panel 136B. The front acoustic treatment panel 136A is located in front of the stage of pre-swirl inlet guide vanes 100, wherein the rear end 138 of the front acoustic treatment panel 136A is located in front of the leading edge 108 of each pre-swirl inlet guide vane 100 of the stage of pre-swirl inlet guide vanes 100. The rear acoustic treatment panel 136B is located in the rear of the stage of pre-swirl inlet guide vanes 100, wherein the front end 140 of the rear acoustic treatment panel 136B is located behind the trailing edge 110 of each pre-swirl inlet guide vane 100 of the stage of pre-swirl inlet guide vanes 100.

[0113] It is worth noting that with this configuration, the length L of the acoustic treatment portion 120 along the axial direction A is AT is along the axial direction A from the front end of the front acoustic treatment panel 136A (at Figure 12 The front end 132 of the acoustic treatment portion 120 is marked in FIG) to the rear end of the last acoustic treatment panel 136B (in FIG). Figure 12 The length L of the rear end 134 of the acoustic treatment portion 120 is marked in FIG. AT .

[0114] Now refer to Figure 13, the leading end 132 of the acoustic treatment portion 120 is positioned at a stage of the pre-swirl inlet guide vanes 100 along the axial direction A. In this manner, it can be understood that, in this context, the term “at” refers to having at least a portion positioned at a position between the leading edge 108 and the trailing edge 110 of each pre-swirl inlet guide vane 100 of the stage of the pre-swirl inlet guide vanes 100 along the axial direction A, or aligned with one of the leading edge 108 and the trailing edge 110 of each pre-swirl inlet guide vane 100 of the stage of the pre-swirl inlet guide vanes 100. More specifically, for the illustrated embodiment, the leading end 132 is positioned at a position between the leading edge 108 and the trailing edge 110 of each pre-swirl inlet guide vane 100 of the stage of the pre-swirl inlet guide vanes 100 along the axial direction A, and the trailing end 134 of the acoustic treatment portion 120 is positioned behind the stage of the pre-swirl inlet guide vanes 100.

[0115] Now refer to Figure 14 The rear end 134 of the acoustic treatment portion 120 is positioned at the stage of the pre-swirl inlet guide vanes 100 along the axial direction A, and more specifically, is positioned between the leading edge 108 and the trailing edge 110 of each pre-swirl inlet guide vane 100 of the stage of the pre-swirl inlet guide vanes 100 along the axial direction A. The front end 132 of the acoustic treatment portion 120 is positioned in front of the stage of the pre-swirl inlet guide vanes 100.

[0116] In short, for Figures 10 to 14 In an exemplary embodiment, the acoustic treatment portion 120, and more specifically, each acoustic treatment panel 136 of the acoustic treatment portion 120, generally includes a linear front end 132 and a linear rear end edge (e.g., the front end 132 is positioned at a common axial position along the circumference of the inner wall 52 of the outer nacelle 50, and similarly, the rear end 134 is positioned at a common axial position along the circumference of the inner wall 52 of the outer nacelle 50).

[0117] Now refer to Figure 15 The acoustic treatment portion 120 includes a rear acoustic treatment panel 136B located behind the stage of pre-swirl inlet guide vanes 100, and a plurality of inter-vane acoustic treatment panels 136C. Each of the plurality of inter-vane acoustic treatment panels 136C is located between adjacent pre-swirl inlet guide vanes 100 of the stage of pre-swirl inlet guide vanes 100. In this manner, the plurality of inter-vane acoustic treatment panels 136C are spaced apart along the circumferential direction C.

[0118] The aft end 142 of each inter-vane acoustic treatment panel 136C meets the forward end 144 of the aft acoustic treatment panel 136B. The forward end 146 of each inter-vane acoustic treatment panel 136C is positioned aft of the leading edge 108 of each pre-swirl inlet guide vane 100 of the stage of pre-swirl inlet guide vanes 100. Furthermore, for the embodiment shown, the forward end 146 of each inter-vane acoustic treatment panel 136C defines a circular shape between adjacent pre-swirl inlet guide vanes 100.

[0119] It is worth noting that with this configuration, the length L of the acoustic treatment portion 120 along the axial direction A is AT The distance from the front end 146 of the inter-blade acoustic treatment panel 136C to the rear end of the rear acoustic treatment panel 136B along the axial direction A is Figure 15 The length L of the rear end 134 of the acoustic treatment portion 120 is marked in FIG. AT .

[0120] In contrast, now refer to Figure 16 The acoustic treatment portion 120 includes a first forward acoustic treatment panel 136A located forward of the pre-swirl stage on the guide vanes, and a plurality of inter-vane acoustic treatment panels 136C. Each of the plurality of inter-vane acoustic treatment panels 136C is located between adjacent pre-swirl inlet guide vanes 100 of a stage of pre-swirl inlet guide vanes 100. A forward end 146 of each inter-vane acoustic treatment panel 136C intersects the forward acoustic treatment panel 136A. A rearward end 142 of each inter-vane acoustic treatment panel 136C is positioned forward of the trailing edge 110 of each pre-swirl inlet guide vane 100 of the stage of pre-swirl inlet guide vanes 100. Furthermore, for the illustrated embodiment, the rearward end 142 of each inter-vane acoustic treatment panel 136C defines a circular shape between adjacent pre-swirl inlet guide vanes 100.

[0121] It is worth noting that with this configuration, the length L of the acoustic treatment portion 120 along the axial direction A is AT is along the axial direction A from the front end of the front acoustic treatment panel 136A (at Figure 16 The length L is the length from the front end 132 of the acoustic treatment portion 120 to the rear end 142 of the inter-blade acoustic treatment panel 136C. AT .

[0122] Now refer to Figure 17 The acoustic treatment portion 120 includes a plurality of acoustic treatment panels 136 spaced apart along the circumferential direction C of the turbofan engine 10. Each of the plurality of acoustic treatment panels 136 defines a circular front end (at Figure 17 The front end 132 of the acoustic treatment portion 120 is marked in FIG) and the rounded rear end (in FIG). Figure 17134 of the acoustic treatment portion 120). For the illustrated embodiment, the leading end of each acoustic treatment panel 136 is located forward of the stage of the pre-swirl inlet guide vanes 100, and the trailing end of each acoustic treatment panel 136 is located at the stage of the pre-swirl inlet guide vanes 100. More specifically, the trailing end of each acoustic treatment panel is located rearward of the leading edge 108 of each pre-swirl inlet guide vane 100 and forward of the trailing edge 110 of each pre-swirl inlet guide vane 100.

[0123] Each acoustic treatment panel 136 generally defines a centerline 148 extending from a front end to a rear end. In the illustrated embodiment, the centerline 148 extends parallel to the axial direction A.

[0124] Now refer to Figure 18 , the acoustic treatment portion 120 again includes a plurality of acoustic treatment panels 136 spaced apart along the circumferential direction C. Figure 18 The acoustic treatment panels 136 each include a front end (at the front of the stage of the pre-swirl inlet guide vane 100) Figure 18 The front end 132 of the acoustic treatment portion 120 is marked in FIG) and the circular rear end of the stage of the pre-swirl inlet guide vane 100 is marked in FIG. Figure 18 Specifically, for the embodiment shown, the rear end of each acoustic treatment panel 136 is aligned with the trailing edge 110 of each pre-swirl inlet guide vane 100 along the axial direction A.

[0125] In the illustrated embodiment, each acoustic treatment panel 136 defines a centerline 148 extending from the front end 132 to the rear end 134. The centerline 148 defines an angle 150 with the axial direction A that is greater than 0 degrees and less than 90 degrees, such as greater than 15 degrees and less than 45 degrees.

[0126] Now refer to Figure 19 , the acoustic processing unit 120 is used with Figure 18 The exemplary embodiments of the present invention are configured in a similar manner. However, for Figure 19 In the embodiment, the front end of each acoustic treatment panel 136 (at Figure 19 The front end 132 of the acoustic treatment portion 120 is positioned at the stage of the pre-swirl inlet guide vane 100, and the rear end of each acoustic treatment panel 136 (at Figure 19 The rear end 134 (marked as the acoustic treatment portion 120 in FIG) is located behind the stage of pre-swirl inlet guide vanes 100 .

[0127] Now refer to Figure 20 , the acoustic processing unit 120 is used with Figure 19 The exemplary acoustic processing section 120 is configured in a similar manner. However, for Figure 20In an embodiment, each acoustic treatment panel 136 defines a rectangular shape as opposed to a generally circular shape. In this manner, the front end (at the Figure 20 The front end 132 of the acoustic treatment portion 120 in FIG. 1 extends in a linear direction perpendicular to the centerline 148 of the corresponding acoustic treatment panel 136. Similarly, the rear end of each acoustic treatment panel 136 (at Figure 20 The rear end 134 of the acoustic treatment portion 120 (marked in FIG) extends in a linear direction perpendicular to the center line 148 of the corresponding acoustic treatment panel 136.

[0128] Now refer to Figure 21 The acoustic treatment portion 120 again includes a plurality of acoustic treatment panels 136, wherein each of the plurality of acoustic treatment panels 136 defines a front end having a circular shape (at Figure 21 The front end 132 of the acoustic treatment portion 120 is marked in FIG) and the rear end having a circular shape (in FIG). Figure 21 The rear end 134 of the acoustic processing unit 120 is marked in FIG. Figure 21 In the embodiment, each acoustic treatment panel 136 is positioned at a stage of the pre-swirl inlet guide vane 100, and more specifically, a front end of each acoustic treatment panel 136 is located at or behind the leading edge 108 of each pre-swirl inlet guide vane 100, and a rear end of each acoustic treatment panel 136 is located at or in front of the trailing edge of each pre-swirl inlet guide vane 100.

[0129] Now refer to Figure 22 , the acoustic processing unit 120 is used with Figure 21 The exemplary embodiments of the present invention are configured in a similar manner. However, for Figure 22 In an embodiment of the present invention, each acoustic treatment panel 136 of the plurality of acoustic treatment panels 136 is at least partially positioned forward of a stage of pre-swirl inlet guide vanes 100 .

[0130] Now refer to Figure 23 The acoustic treatment portion 120 again includes a plurality of acoustic treatment panels 136 spaced apart along the circumferential direction C of the turbofan engine 10, wherein each acoustic treatment panel 136 includes a front end defining a circular shape (at Figure 23 The front end 132 of the acoustic treatment portion 120 is marked in FIG) and the rear end defining a circular shape (in FIG). Figure 23 The rear end 134 of the acoustic processing unit 120 is marked in FIG. Figure 23 In the embodiment, the front end of each acoustic treatment panel 136 is located in front of the stage of pre-swirl inlet guide vanes 100 , and the rear end of each acoustic treatment panel 136 is located behind the stage of pre-swirl inlet guide vanes 100 .

[0131] In addition, for Figure 23In the embodiment of FIG. 1 , each acoustic treatment panel 136 defines a centerline 148. However, Figure 23 The centerline 148 of each acoustic treatment panel 136 in the embodiment of FIG extends at least partially in a nonlinear direction. For example, in FIG. Figure 23 In the embodiment of the present invention, the centerline 148 of each acoustic treatment panel 136 defines a curve, wherein the centerline 148 extends through the stage of the pre-swirl inlet guide vanes 100. The curve of the centerline 148 of each acoustic treatment panel 136 is generally complementary to the camber line 152 of the pre-swirl inlet guide vanes 100.

[0132] Further aspects are provided by the subject matter of the following clauses:

[0133] A gas turbine engine defining an axial direction and a radial direction, the gas turbine engine comprising: a turbine; a fan rotatable by the turbine, the fan comprising fan blades defining an outer tip along the radial direction, a trailing edge at the tip, and a length L along the axial direction at the tip FB and an outer nacelle surrounding the fan and at least partially surrounding the turbine, the outer nacelle including a stage of pre-swirl inlet guide vanes located upstream of the fan, the stage of pre-swirl inlet guide vanes having pre-swirl inlet guide vanes defining a 15% span position and a leading edge at the 15% span position, the pre-swirl inlet guide vanes further defining a length L along the axial direction at the 15% span position. IGV The outer nacelle further includes an inner surface along the radial direction and an acoustic treatment portion coupled to or integrated with the inner surface, the acoustic treatment portion defining a length L along the axial direction. AT ; wherein the gas turbine engine defines a length L from the leading edge of the pre-swirl inlet guide vane at the 15% span position to the trailing edge of the fan blade at the tip of the fan blade S , and wherein the length L of the acoustic treatment portion AT As follows: (L IGV 2 / L S )×UCF1 <L AT <(L FB 3 / L IGV )×UCF2, where UCF1 is equal to 1 -1 The first unit correction factor of inches, and UCF2 is equal to 1 -2 Second unit correction factor for inches.

[0134] A gas turbine engine according to any preceding clause, wherein the length L IGV greater than or equal to 0.3 inches and less than or equal to 11 inches, and wherein the length L FB Greater than or equal to 3 inches and less than or equal to 15 inches.

[0135] A gas turbine engine according to any preceding clause, wherein the length L IGV greater than or equal to 1.5 inches and less than or equal to 9 inches, and wherein the length L FB Greater than or equal to 5 inches and less than or equal to 11 inches.

[0136] A gas turbine engine according to any preceding clause, wherein the length L S Greater than or equal to 4 inches and less than or equal to 30 inches.

[0137] A gas turbine engine according to any preceding clause, wherein the length L S Greater than or equal to 7.5 inches and less than or equal to 24 inches.

[0138] The gas turbine engine of any preceding clause, wherein the acoustic treatment is integrated into the inner surface of the outer nacelle.

[0139] The gas turbine engine of any preceding clause, wherein the acoustic treatment is positioned at least partially along the axial direction between the stage of the pre-swirl inlet guide vanes and the tip of the fan blade.

[0140] A gas turbine engine according to any preceding clause, wherein the acoustic treatment is positioned at least partially along the axial direction at a stage of the pre-swirl inlet guide vanes.

[0141] The gas turbine engine of any preceding clause, wherein the acoustic treatment is located at least partially along the axial direction at the tip of the fan blade.

[0142] A gas turbine engine as claimed in any preceding clause, wherein the acoustic treatment is at least partially integrated with a surface of the pre-swirl inlet guide vanes.

[0143] The gas turbine engine of any preceding clause, wherein the acoustic treatment extends continuously along the circumference of the inner surface of the outer nacelle.

[0144] A gas turbine engine as claimed in any preceding clause, wherein the acoustic treatment comprises a plurality of acoustic treatment panels.

[0145] The gas turbine engine of any preceding clause, wherein the plurality of acoustic treatment panels are spaced apart along a circumferential direction of the gas turbine engine.

[0146] The gas turbine engine of any preceding clause, wherein the plurality of acoustic treatment panels are spaced apart along the axial direction of the gas turbine engine.

[0147] A gas turbine engine as claimed in any preceding clause, wherein the acoustic treatment extends from a position forward of the stage of pre-swirl inlet guide vanes to a position aft of the stage of pre-swirl inlet guide vanes.

[0148] A gas turbine engine according to any preceding clause, wherein the gas turbine engine is configured as a direct drive gas turbine engine.

[0149] A gas turbine engine according to any preceding clause, wherein the pre-swirl inlet guide vanes define a span, wherein the fan blades define a fan diameter, and wherein the span is greater than or equal to 5% and less than or equal to 50% of the fan blade diameter.

[0150] A gas turbine engine according to any preceding clause, wherein the pre-swirl inlet guide vanes are attached to or integrated with a main body of the nacelle.

[0151] A gas turbine engine as claimed in any preceding clause, wherein the acoustic treatment comprises a perforated plate and a hollow body.

[0152] A gas turbine engine as in any preceding clause, wherein the acoustic treatment comprises at least a portion extending to the pre-swirl inlet guide vanes.

[0153] This written description uses examples to disclose the present disclosure, including the best mode, and also to enable any person skilled in the art to practice the present disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the present disclosure 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. A gas turbine engine defining an axial direction and a radial direction, characterized in that: The gas turbine engine comprises: turbines; a fan rotatable by the turbine, the fan comprising fan blades defining an outer tip in the radial direction, a trailing edge at the outer tip, and a length L in the axial direction at the outer tip FB ;and an outer nacelle surrounding the fan and at least partially surrounding the turbine, the outer nacelle including a stage of pre-swirl inlet guide vanes located upstream of the fan, the stage of pre-swirl inlet guide vanes having pre-swirl inlet guide vanes, the pre-swirl inlet guide vanes defining a 15% span position and a leading edge at the 15% span position, the pre-swirl inlet guide vanes further defining a length L along the axial direction at the 15% span position IGV The outer nacelle further includes an inner surface along the radial direction and an acoustic treatment portion coupled to or integrated with the inner surface, the acoustic treatment portion defining a length L along the axial direction. AT ; wherein the gas turbine engine defines a length L from the leading edge of the pre-swirl inlet guide vane at the 15% span position to the trailing edge of the fan blade at the outer tip of the fan blade S ,and The length L of the acoustic treatment portion AT As follows: (L IGV 2 / L S )×UCF1 <L AT <(L FB 3 / L IGV )×UCF2, Where UCF1 is equal to 1 -1 The first unit correction factor of inches, and UCF2 is equal to 1 -2 Second unit correction factor for inches.

2. The gas turbine engine according to claim 1, wherein: Wherein the length L IGV greater than or equal to 0.3 inches and less than or equal to 11 inches, and wherein the length L FB Greater than or equal to 3 inches and less than or equal to 15 inches.

3. The gas turbine engine according to claim 1, wherein: Wherein the length L IGV greater than or equal to 1.5 inches and less than or equal to 9 inches, and wherein the length L FB Greater than or equal to 5 inches and less than or equal to 11 inches.

4. The gas turbine engine according to claim 1, wherein: Wherein the length L S Greater than or equal to 4 inches and less than or equal to 30 inches.

5. The gas turbine engine according to claim 1, wherein Wherein the length L S Greater than or equal to 7.5 inches and less than or equal to 24 inches.

6. The gas turbine engine according to claim 1, wherein: wherein the acoustic treatment is integrated into the inner surface of the outer nacelle.

7. The gas turbine engine according to claim 1, wherein: The acoustic treatment portion is at least partially positioned along the axial direction between the stage of the pre-swirl inlet guide vane and the outer tip of the fan blade.

8. The gas turbine engine according to claim 1, wherein: The acoustic treatment portion is at least partially positioned at a stage of the pre-swirl inlet guide vanes along the axial direction.

9. The gas turbine engine according to claim 1, wherein: Wherein the acoustic treatment portion is at least partially positioned at the outer tip of the fan blade along the axial direction.

10. The gas turbine engine according to claim 1, wherein: The acoustic treatment portion is at least partially integrated with the surface of the pre-swirl inlet guide vane.

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