Ice accretion mechanism for a turbofan engine

Abstract

A turbofan engine is provided. The turbofan engine includes a fan including a plurality of fan blades, a turbine operably coupled to the fan for driving the fan, the turbine including a compressor section, a combustion section, and a turbine section in serial flow order and together defining a core air flowpath, an nacelle surrounding and at least partially enclosing the fan, an inlet pre-swirl feature upstream of the plurality of fan blades, the inlet pre-swirl feature attached to or integrated into the nacelle, and a device on the inlet pre-swirl feature for reducing ice accumulation or ice formation, the device in communication with the inlet pre-swirl feature.

Description

TECHNICAL FIELD

[0001] The present subject matter relates generally to a gas turbine engine, or more specifically, to a gas turbine engine configured to reduce ice accumulation or ice formation on inlet components of the engine. BACKGROUND

[0002] A turbofan engine generally includes a fan and a turbine arranged in flow communication with one another, the fan having a plurality of fan blades. Additionally, the turbine of the turbofan engine generally includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. In operation, air is provided from the fan to an inlet of the compressor section where one or more axial compressors progressively compress the air until it reaches the combustion section. Fuel is mixed with the compressed air within the combustion section and combusted 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, e.g., to the atmosphere.

[0003] However, during inclement weather, freezing rain, hail, sleet, ice, etc. can accumulate on the inlet components of the turbofan engine. When ice accumulates, it can break off and be ingested into the engine. Further, much of the ice can damage the fan blades or other downstream components of the engine and can cause the engine to flame out. BRIEF DESCRIPTION OF DRAWINGS

[0004] In the description of the specification, the complete disclosure of the present disclosure including its best mode is set forth, as it would be to one of ordinary skill in the art, in which:

[0005] Figure 1 is a schematic cross-sectional view of an exemplary gas turbine engine in accordance with an exemplary embodiment of the present subject matter.

[0006] Figure 2 is a schematic cross-sectional view of an exemplary gas turbine engine in accordance with an exemplary embodiment of the present subject matter. Figure 1 is a close-up schematic cross-sectional view of a forward end of the exemplary gas turbine engine of

[0007] Figure 3 is a view along an axial direction of the gas turbine engine of Figure 1 Figure 1 is a schematic view of an inlet of the exemplary gas turbine engine of

[0008] Figure 4 is a schematic view of an inlet of a gas turbine engine in accordance with another exemplary embodiment of the present disclosure.

[0009] Figure 5A ​is a schematic cross-sectional view of an exemplary gas turbine engine according to an exemplary embodiment of the present subject matter.

[0010] Figure 5B is a cross-sectional view of a partial span inlet guide vane of an exemplary gas turbine engine according to another exemplary embodiment of the present subject matter at a first location along a span of the partial span inlet guide vane.

[0011] Figure 6 is a schematic cross-sectional view of an exemplary gas turbine engine according to another exemplary embodiment of the present subject matter.

[0012] Figure 7 is a schematic cross-sectional view of an exemplary gas turbine engine according to another exemplary embodiment of the present subject matter.

[0013] Figure 8A is a schematic cross-sectional view of an exemplary gas turbine engine according to another exemplary embodiment of the present subject matter.

[0014] Figure 8B is a cross-sectional view of a partial span inlet guide vane of an exemplary gas turbine engine according to another exemplary embodiment of the present subject matter at a first location along a span of the partial span inlet guide vane.

[0015] Figure 9A is a schematic cross-sectional view of an exemplary gas turbine engine according to another exemplary embodiment of the present subject matter.

[0016] Figure 9B is a schematic cross-sectional view of an exemplary gas turbine engine according to another exemplary embodiment of the present subject matter. Figure 9A is a cross-sectional view of a partial span inlet guide vane of an exemplary gas turbine engine according to another exemplary embodiment of the present subject matter at a first location along a span of the partial span inlet guide vane.

[0017] Corresponding reference symbols have been used throughout the various drawings and specific reference characters indicate corresponding parts among the figures. The examples listed herein illustrate exemplary embodiments of the present disclosure, and these examples should not be construed as limiting the scope of the present disclosure in any way. DETAILED DESCRIPTION

[0018] Reference will now be made in detail to embodiments of the present disclosure, one or more examples of which are illustrated in the drawings. The detailed description uses numerical and letter designations

[0019] The following description is provided so as to enable any person skilled in the art to make and use the described embodiments of the disclosure. Various modifications, equivalents, changes, and substitutions of the described embodiments are possible. Any and all such modifications, variations, permutations, and equivalents are intended to fall within the scope of the present disclosure.

[0020] The word "exemplary" is used herein to mean "serving as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other implementations. Furthermore, unless otherwise explicitly provided, none of the embodiments described herein are intended to be interpreted as favoring one implementation of the disclosure over another.

[0021] For purposes of the description hereinafter, the terms "upper", "lower", "right", "left", "vertical", "horizontal", "top", "bottom", "lateral", "longitudinal", and derivatives thereof shall relate to the disclosure as it is oriented in the drawing

[0022] As used herein, the terms "first", "second", and "third" can be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components.

[0023] The terms "forward" and "aft" refer to relative positions within a gas turbine engine, with forward referring to a position closer to an engine inlet and aft referring to a position closer to an engine nozzle or exhaust.

[0024] The terms "upstream" and "downstream" refer to the relative direction with respect to fluid flow in a fluid pathway. For example, "upstream" refers to the direction from which the fluid flows, and "downstream" refers to the direction to which the fluid flows.

[0025] The singular forms "a", "an", and "the" include plural referents unless the context clearly dictates otherwise.

[0026] Further, unless otherwise specified, the terms "low," "high," or their respective comparative forms (e.g., lower, higher, as applicable) each refer to relative speeds and pressures within the engine. For example, a "low pressure turbine" operates at generally lower pressures than a "high pressure turbine." Alternatively, unless otherwise specified, the aforementioned terms can be interpreted as superlative. For example, a "low pressure turbine" can refer to the lowest maximum pressure turbine within a turbine section, and a "high pressure turbine" can refer to the highest maximum pressure turbine within a turbine section. The engines of the present disclosure can also include an intermediate pressure turbine, such as an engine having three spools.

[0027] Approximating language as used herein with respect to a quantitative representation of a parameter is applied to modify any quantitative representation that could permissibly vary from the stated limitation without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms such as "about," "approximately,” and "substantially” are not limited to the precise value specified. In at least some instances, the approximating language can correspond to the precision of an instrument for measuring the value or the precision of the method or machine for constructing or manufacturing the component and / or system. For example, the approximating language can refer to a margin of error within 1%, 2%, 4%, 10%, 15%, or 20% of the stated value. These approximating margins can apply to individual values, to any one of the endpoints of a range, or to the range of endpoints as a whole.

[0028] Ratios, in this and other contexts, are combined and / or interchanged, unless otherwise indicated or unless context clearly dictates to the contrary, such ranges are identified and include all subranges therein. For instance, all ranges disclosed herein are inclusive of the endpoints, and endpoints are combinable with one another to form ranges.

[0029] As used herein, the term "fan pressure ratio" refers to the ratio of the air pressure immediately downstream of the fan blades of a fan during operation of the fan to the air pressure immediately upstream of the fan blades of the fan during operation of the fan.

[0030] As used herein, the term "rated speed" with respect to a turbofan engine refers to the maximum rotational speed that the turbofan engine can achieve during normal operation. For example, a turbofan engine can operate at rated speed during maximum load operation, such as during takeoff operation.

[0031] Also as used herein, the term "fan tip speed" defined by a plurality of fan blades of a fan refers to the linear speed of the outer tips of the fan blades in a radial direction during operation of the fan.

[0032] The present disclosure generally relates to devices for reducing ice accumulation or ice formation on inlet components of an engine, such as inlet pre-swirl features configured as multiple span inlet guide vanes.

[0033] In some example embodiments of the present disclosure, the apparatus for reducing ice accumulation or ice formation includes a heat source in thermal communication with the inlet pre-swirl feature, e.g., the inlet pre-swirl feature is configured as a plurality of partial span inlet guide vanes. For example, in a first example embodiment, the heat source includes an engine exhaust gas stream. In another example embodiment, the heat source includes an electric heating element disposed in a leading edge of the inlet pre-swirl feature. In yet another example embodiment, the heat source includes engine oil.

[0034] In other example embodiments, the apparatus for reducing ice accumulation or ice formation includes a vibration assembly in mechanical communication with the inlet pre-swirl feature or an anti-icing coating on the pre-swirl feature, e.g., the inlet pre-swirl feature is configured as a plurality of partial span inlet guide vanes.

[0035] The inclusion of one or more of these apparatus for reducing ice accumulation or ice formation provides an anti-icing or de-icing mechanism that prevents ice buildup and shedding into the engine during, for example, adverse weather conditions, resulting in safer operation of the gas turbine engine.

[0036] Reference will now be made to the drawings wherein like numerals refer to like components throughout. Figure 1 is a schematic cross-sectional view of a gas turbine engine in accordance with example embodiments of the present disclosure. More specifically, for Figure 1 embodiments of the present disclosure, the gas turbine engine is an aircraft turbofan jet engine 10 (referred to herein as "turbofan engine 10") configured to be mounted to an aircraft, such as in a wing- under configuration or a tail-mounted configuration. As shown, Figure 1 the turbofan engine 10 defines an axial direction A (extending parallel to a longitudinal centerline 12 provided for reference), a radial direction R, and a circumferential direction (i.e., a direction extending about the axial direction A; not shown). Generally, the turbofan engine 10 includes a fan section 14 and a turbine machine 16 (sometimes also referred to or alternatively as a "core turbine engine") disposed downstream from the fan section 14.

[0037] The depicted exemplary turbine 16 generally includes a substantially tubular outer casing 18 defining an annular inlet 20. The casing 18 encloses, in serial flow relationship: a compressor section including a first booster or low pressure (LP) compressor 22 and a second high pressure (HP) compressor 24; a combustion section 26; a turbine section including a first high pressure (HP) turbine 28 and a second low pressure (LP) turbine 30; and an ejection exhaust nozzle section 32. A high pressure (HP) shaft 34 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, the combustion section 26, the turbine section, and the ejection exhaust nozzle section 32 are arranged in serial 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 a mid-pressure turbine (e.g., engines having three spools).

[0038] Still referring to Figure 1 embodiments of the present disclosure, the fan section 14 includes a variable pitch single stage fan 38, the turbine 16 being operably coupled to the fan 38 to drive the fan 38. The fan 38 includes a plurality of rotatable fan blades 40 coupled to a disk 42 in a spaced apart manner. As shown, the fan blades 40 extend outwardly from the disk 42 in a generally radial direction R. Each fan blade 40 is rotatable about a pitch axis P relative to the disk 42 by virtue of the fan blades 40 being operably coupled to a suitable actuation member 44, the actuation member 44 being configured to, for example, uniformly collectively change the pitch of the fan blades 40. The fan blades 40, the disk 42, and the actuation member 44 are rotatable together about the longitudinal centerline 12 by the LP shaft 36 across a power gear box 46. The power gear box 46 includes a plurality of gears for reducing the rotational speed of the LP shaft 36 to a more efficient rotational fan speed. Thus, for the depicted embodiment, the turbine 16 is operably coupled to the fan 38 by the power gear box 46.

[0039] In exemplary embodiments, the fan section 14 includes twenty-two (22) or fewer fan blades 40. In certain exemplary embodiments, the fan section 14 includes twenty (20) or fewer fan blades 40. In certain exemplary embodiments, the fan section 14 includes eighteen (18) or fewer fan blades 40. In certain exemplary embodiments, the fan section 14 includes sixteen (16) or fewer fan blades 40. In certain exemplary embodiments, it is contemplated that the fan section 14 includes other numbers of fan blades 40 for particular applications.

[0040] During operation of the turbofan engine 10, the fan 38 defines a fan pressure ratio, and the plurality of fan blades 40 each define a fan tip speed. The exemplary turbofan engine 10 shown defines a relatively high fan tip speed and a relatively low fan pressure ratio. As used herein, the term "fan pressure ratio" refers to the ratio of the air pressure immediately downstream of the fan blades 40 during operation of the fan 38 to the air pressure immediately upstream of the fan blades 40 during operation of the fan 38. For example, the fan pressure ratio of the turbofan engine 10 can be less than or equal to about 1.5. For example, in certain exemplary embodiments, the turbofan engine 10 can define a fan pressure ratio of less than or equal to about 1.4. The fan pressure ratio can be the fan pressure ratio of the fan 38 during operation of the turbofan engine 10, such as during operation of the turbofan engine 10 at a rated speed. Figure 1 For the embodiment shown, the fan 38 of the turbofan engine 10 defines a relatively low fan pressure ratio. For example, the depicted turbofan engine 10 defines a fan pressure ratio of less than or equal to about 1.5. For example, in certain exemplary embodiments, the turbofan engine 10 can define a fan pressure ratio of less than or equal to about 1.4. The fan pressure ratio can be the fan pressure ratio of the fan 38 during operation of the turbofan engine 10, such as during operation of the turbofan engine 10 at a rated speed.

[0041] As used herein with respect to the turbofan engine 10, the term "rated speed" refers to the maximum rotational speed that the turbofan engine 10 can achieve while in normal operation. For example, the turbofan engine 10 can operate at a rated speed during maximum load operation, such as during takeoff operation.

[0042] Also as used herein, the term "fan tip speed" defined by the plurality of fan blades 40 refers to the linear speed of the outer tip of the fan blades 40 in the radial direction R during operation of the fan 38. In exemplary embodiments, the turbofan engine 10 of the present disclosure rotates the fan blades 40 of the fan 38 at a relatively high rotational speed. For example, during operation of the turbofan engine 10 at a rated speed, the fan tip speed of each of the plurality of fan blades 40 is greater than or equal to 1000 feet per second and less than or equal to 2250 feet per second. In certain exemplary embodiments, during operation of the turbofan engine 10 at a rated speed, the fan tip speed of each fan blade 40 can be greater than or equal to 1,250 feet per second and less than or equal to 2250 feet per second. In certain exemplary embodiments, during operation of the turbofan engine 10 at a rated speed, the fan tip speed of each fan blade 40 can be greater than or equal to about 1,350 feet per second, such as greater than about 1,450 feet per second, such as greater than about 1,550 feet per second, and less than or equal to 2250 feet per second.

[0043] Still referring to Figure 1In exemplary embodiments, the disk 42 is covered by a rotatable front nacelle or hub 48 that is aerodynamically shaped to facilitate airflow through the plurality of fan blades 40. Additionally, the exemplary fan section 14 includes an annular fan casing or outer nacelle 50 that at least partially and, for the depicted embodiment, circumferentially surrounds at least a portion of the fan 38 and turbine 16.

[0044] More particularly, the outer nacelle 50 includes an inner wall 52, and a downstream section 54 of the inner wall 52 of the outer nacelle 50 extends over an outer portion of the turbine 16 so as to define a bypass airflow passage 56 therebetween. Further, for the depicted embodiment, the outer nacelle 50 is supported relative to the turbine 16 by a plurality of circumferentially spaced apart outlet guide vanes 55.

[0045] During operation of the turbofan engine 10, an amount of air 58 enters the turbofan engine 10 through an associated inlet 60 of the outer nacelle 50 and / or fan section 14. As the amount of air 58 passes through the fan blades 40, a first portion of the air 58 is directed or channeled into the bypass airflow passage 56 as indicated by arrow 62, and a second portion of the air 58 is directed or channeled into the core airflow path 37 as indicated by arrow 64. The ratio between the amount of airflow through the bypass airflow passage 56 (i.e., the first portion of air indicated by arrow 62) and the amount of airflow through the core airflow path 37 (i.e., the second portion of air indicated by arrow 64) is referred to as the bypass ratio.

[0046] In exemplary embodiments, the bypass ratio during operation of the turbofan engine 10 (e.g., at rated speed) is less than or equal to about eleven (11). For example, the bypass ratio during operation of the turbofan engine 10 (e.g., at rated speed) can be less than or equal to about ten (10), such as less than or equal to about nine (9). Additionally, the bypass ratio can be at least about two (2).

[0047] In other exemplary embodiments, the bypass ratio can be substantially between about 7: 1 and about 20: 1, such as between about 10: 1 and about 18: 1. The second portion of air indicated by arrow 64 is then channeled through the high pressure (HP) compressor 24 and into the combustion section 26, where the pressure of the second portion of air increases, the second portion of air is mixed with fuel and combusted to provide combustion gases 66 at the combustion section 26.

[0048] In exemplary embodiments, the gear ratio of the power gear box 46 is greater than or equal to 1.2 and less than or equal to 3.0. In some exemplary embodiments, the gear ratio of the power gear box 46 is greater than or equal to 1.2 and less than or equal to 2.6. In other exemplary embodiments, the gear ratio of the power gear box 46 is greater than or equal to 1.2 and less than or equal to 2.0.

[0049] Further, as described herein, the turbofan engine of the present disclosure also provides a pre-swirl flow ahead of the fan blade tip.

[0050] Still referring to Figure 1 A second portion of the compressed air from the compressor section, indicated by arrow 64, is mixed with fuel and combusted within the combustion section to provide combustion gases 66. The combustion gases 66 are directed from the combustion section 26 through the HP turbine 28, a portion of the thermal and / or kinetic energy from the combustion gases 66 being extracted at the HP turbine 28 via successive stages of HP turbine stator vanes 68 coupled to the outer casing 18 and HP turbine rotor blades 70 coupled to the HP shaft 34, thus rotating the HP shaft 34 to thereby support operation of the HP compressor 24. The combustion gases 66 are then directed through the LP turbine 30, a second portion of the thermal and kinetic energy being extracted from the combustion gases 66 at the LP turbine 30 via successive stages of LP turbine stator vanes 72 coupled to the outer casing 18 and LP turbine rotor blades 74 coupled to the LP shaft 36, thus rotating the LP shaft 36 to thereby support operation of the LP compressor 22 and / or rotation of the fan 38.

[0051] Subsequently, the combustion gases 66 are directed through the ejection exhaust nozzle section 32 of the turbine engine 16 to provide propulsive thrust. At the same time, 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, the pressure of the first portion of air, indicated by arrow 62, is significantly increased, also providing propulsive thrust. The HP turbine 28, the LP turbine 30, and the ejection exhaust nozzle section 32 at least partially define a hot gas path 78 for directing the combustion gases 66 through the turbine engine 16.

[0052] Still referring to Figure 1 The turbofan engine 10 additionally includes a device 200 for reducing ice accumulation or ice formation on the inlet pre-swirl feature, e.g., the inlet pre-swirl feature is configured as a plurality of partial span inlet guide vanes 100, as described in greater detail herein.

[0053] In some example embodiments, it will be appreciated that the example turbofan engine 10 of the present disclosure can be a relatively large power class turbofan engine 10. Thus, when operating at a rated speed, the turbofan engine 10 can be configured to generate a relatively large amount of thrust. More particularly, when operating at a rated speed, the turbofan engine 10 can be configured to generate at least about 20,000 pounds of thrust, such as at least about 25,000 pounds of thrust, such as at least about 30,000 pounds of thrust, and up to, for example, about 150,000 pounds of thrust. Thus, the turbofan engine 10 can be referred to as a relatively large power class gas turbine engine.

[0054] Further, it will be appreciated that,Figure 1 The exemplary turbofan engine 10 depicted is by way of example only, and in other exemplary embodiments, the turbofan engine 10 can have any other suitable configuration. For example, in certain exemplary embodiments, the fan can not be a variable pitch fan, the engine can not include a reduction gearbox (e.g., power gear box 46) to drive the fan, can include any other suitable number or arrangement of shafts, spools, compressors, turbines, etc.

[0055] As discussed above, the turbofan engine 10 of the present disclosure also provides a pre-swirl flow in front of the tips of the fan blades 40. Reference is now also made to Figure 2 A close-up cross-sectional view of the fan section 14 and the forward end of the turbine 16 of the exemplary turbofan engine 10 of Figure 1 In exemplary embodiments, the turbofan engine 10 includes an inlet pre-swirl feature that is located upstream of the plurality of fan blades 40 of the fan 38 and is attached to or integrated into the outer nacelle 50. More specifically, for embodiments of Figure 1 and 2 The inlet pre-swirl feature is configured as a plurality of partial-span inlet guide vanes 100. Each of the plurality of partial-span inlet guide vanes 100 cantilevers from the outer nacelle 50 (such as from the inner wall 52 of the outer nacelle 50) at a location that is forward of the plurality of fan blades 40 of the fan 38 and aft of the inlet 60 of the outer nacelle 50 along the axial direction A. More specifically, each of the plurality of partial-span inlet guide vanes 100 defines an outer end 102 along the radial direction R and is attached / connected to the outer nacelle 50 at the outer end 102 by suitable connecting means (not shown). For example, each of the plurality of partial-span inlet guide vanes 100 can be bolted to the inner wall 52 of the outer nacelle 50 at the outer end 102, welded to the inner wall 52 of the outer nacelle 50 at the outer end 102, or attached to the outer nacelle 50 in any other suitable manner at the outer end 102.

[0056] Further, for the depicted embodiment, the plurality of partial-span inlet guide vanes 100 extend generally in the radial direction R from the outer end 102 to the inner end 104 (i.e., the inner end 104 is along the radial direction R). Moreover, as will be appreciated, for the depicted embodiment, each of the plurality of partial-span inlet guide vanes 100 is unconnected at the respective inner end 104 with an adjacent partial-span inlet guide vane 100 (i.e., the adjacent partial-span inlet guide vanes 100 do not contact one another at the radially inner end 104 and do not include any intermediate connecting members, such as connecting rings, struts, etc., at the radially inner end 104). More specifically, for the depicted embodiment, each partial-span inlet guide vane 100 is fully supported at the respective outer end 102 by connection with the outer nacelle 50 (and not by, for example, any structure extending between adjacent partial-span inlet guide vanes 100 at a location inward of the outer end 102 along the radial direction R). As will be discussed below, this can reduce the amount of turbulence generated by the partial-span inlet guide vanes 100.

[0057] Moreover, as shown, each of the plurality of partial-span inlet guide vanes 100 does not extend entirely between the outer nacelle 50 and, for example, the hub 48 of the turbofan engine 10. More specifically, for the depicted embodiment, each of the plurality of inlet guide vanes defines an inlet guide vane ("IGV") span 106 along the radial direction R, and further, each of the plurality of partial-span inlet guide vanes 100 further defines a leading edge 108 and a trailing edge 110. The IGV span 106 refers to a measure along the radial direction R between the outer end 102 and the inner end 104 of the partial-span inlet guide vane 100 at the leading edge 108 of the partial-span inlet guide vane 100. Similarly, it will be appreciated that 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 the fan 38 further defines a leading edge 114 and a trailing edge 116, and the fan blade span 112 refers to a measure along the radial direction R between a radially outer tip and a base of the respective fan blade 40 at the leading edge 114 of the fan blade 40.

[0058] For the depicted embodiment, the IGV span 106 is at least about five percent of the fan blade span 112, and up to about fifty-five percent of the fan blade span 112. For example, in certain example embodiments, the IGV span 106 can be between about fifteen percent of the fan blade span 112 and about forty-five percent of the fan blade span 112, such as between about thirty percent of the fan blade span 112 and about forty percent of the fan blade span 112.

[0059] Reference will now also be made to Figure 3, providing Figure 1 and 2 an axial view of an inlet 60 of a turbofan engine 10 according to another example embodiment of the disclosure. For the depicted embodiment, the turbofan engine 10 includes less than twenty partial-span inlet guide vanes 100. More specifically, for the depicted embodiment, the turbofan engine 10 includes at least eight partial-span inlet guide vanes 100, or more specifically, includes exactly eight partial-span inlet guide vanes 100. Additionally, for the depicted embodiment, the plurality of partial-span inlet guide vanes 100 are substantially uniformly spaced apart along the circumferential direction C. More specifically, each of the plurality of partial-span inlet guide vanes 100 defines a circumferential spacing 118 with an adjacent partial-span inlet guide vane 100, wherein the circumferential spacing 118 is substantially equal between each adjacent partial-span inlet guide vane 100.

[0060] While not depicted, in certain example embodiments, the number of partial-span inlet guide vanes 100 can 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 partial-span inlet guide vanes 100 can be greater than the number of fan blades 40 of the fan 38 of the turbofan engine 10, or alternatively, can be less than the number of fan blades 40 of the fan 38 of the turbofan engine 10.

[0061] Further, it should be appreciated that in other example embodiments, the turbofan engine 10 can include any other suitable number of partial-span inlet guide vanes 100 and / or circumferential spacing 118 of the partial-span inlet guide vanes 100. For example, now briefly referring to Figure 4 , providing an axial view of an inlet 60 of a turbofan engine 10 according to another example embodiment of the disclosure. For the depicted embodiment, the turbofan engine 10 includes less than twenty partial-span inlet guide vanes 100. More specifically, for the depicted embodiment, the turbofan engine 10 includes at least eight partial-span inlet guide vanes 100, or more specifically, includes exactly eight partial-span inlet guide vanes 100. Additionally, for the depicted embodiment, the plurality of partial-span inlet guide vanes 100 are substantially uniformly spaced apart along the circumferential direction C. More specifically, each of the plurality of partial-span inlet guide vanes 100 defines a circumferential spacing 118 with an adjacent partial-span inlet guide vane 100, wherein the circumferential spacing 118 is substantially equal between each adjacent partial-span inlet guide vane 100. Figure 4 Figure 4 Figure 4 ​​In the depicted embodiment, the plurality of partial-span inlet guide vanes 100 are substantially non-uniformly spaced in the circumferential direction C. For example, at least some of the plurality of partial-span inlet guide vanes 100 define a first circumferential spacing 118A, while others of the plurality of partial-span inlet guide vanes 100 define a second circumferential spacing 118B. For the depicted embodiment, the first circumferential spacing 118A is at least about twenty percent greater than the second circumferential spacing 118B, such as at least about twenty-five percent greater, such as at least about thirty percent greater, such as up to about two hundred percent greater. Notably, as will be described in greater detail below, the circumferential spacing 118 refers to the average circumferential spacing between adjacent partial-span inlet guide vanes 100. The non-uniform circumferential spacing can be, for example, offset upstream of the partial-span inlet guide vanes 100.

[0062] Referring again to Figure 2 It will be appreciated that each of the plurality of partial-span inlet guide vanes 100 is configured to be upstream of the plurality of fan blades 40 of the fan 38 to pre- swirl the airflow 58 provided by the inlet 60 of the nacelle 50. As briefly discussed above, pre-swirling the airflow 58 prior to the airflow 58 reaching the plurality of fan blades 40 of the fan 38 can reduce separation losses and / or impingement losses, thereby allowing the fan 38 to operate at the relatively high fan tip speeds described above with less efficiency losses.

[0063] As discussed, the present disclosure provides a device 200 for reducing ice accumulation or ice formation on an inlet pre-swirl feature, e.g., the inlet pre-swirl feature is configured as a plurality of partial-span inlet guide vanes 100, in communication with the inlet pre-swirl feature. This provides an anti-icing or de-icing mechanism that prevents ice buildup and shedding into the engine during adverse weather conditions.

[0064] Referring now generally to Figures 5A to 8B In the exemplary embodiment of the present disclosure, the device 200 for reducing ice accumulation or ice formation includes a heat source 202 in thermal communication with the inlet pre-swirl feature, e.g., the inlet pre-swirl feature is configured as a plurality of partial-span inlet guide vanes 100.

[0065] Referring to Figure 5A A close-up cross-sectional view of the fan section 14 and forward end of the turbine 16 of the turbofan engine 10 according to the exemplary embodiment of the present disclosure is provided. Figure 5A The exemplary engine 10 can be configured in a similar manner as the exemplary engine described above. Figure 2 The exemplary engine can be configured in a similar manner as the exemplary engine described above.

[0066] In the depicted example embodiment, the heat source 202 includes the engine bleed air stream 160. For example, in a first example embodiment, with reference to Figure 5A The heat source 202 includes an exhaust air supply assembly 162 that includes an exhaust air supply duct 164 in airflow communication with a high pressure air source 165 (e.g., the engine bleed air stream 160 from the engine 10) and a leading edge 108 of an inlet pre-whirl feature, e.g., configured as a plurality of partial span inlet guide vanes 100. The exhaust air supply duct 164 is configured to receive the engine bleed air stream 160 and provide it to the leading edge 108 of the inlet pre-whirl feature, e.g., configured as a plurality of partial span inlet guide vanes 100. For example, the exhaust air supply duct 164 can provide the engine bleed air stream 160 to a location proximate the leading edge 108, e.g., to a location closer to the leading edge 108 than to the trailing edge 110.

[0067] In example embodiments, the high pressure air source 165 is a compressor section of the turbine engine 16, e.g., the LP compressor 22. For example, in example embodiments, the hot compressor discharge air, e.g., the engine bleed air stream 160, is directed to the leading edge 108 of the partial span inlet guide vanes 100, and then the cooler engine bleed air stream 160 is returned to the turbine engine 16. It is contemplated that the compressor discharge air can originate from any stage of the LP compressor 22 or the HP compressor 24.

[0068] In this manner, the engine bleed air stream 160 is used to heat the partial span inlet guide vanes 100 and operates as a device for reducing ice accumulation or ice formation at the inlet pre-whirl feature, e.g., configured as a plurality of partial span inlet guide vanes 100.

[0069] In example embodiments, the exhaust air supply assembly 162 further includes an exhaust air return duct 172 in airflow communication with the exhaust air supply duct 164 and a portion of the turbine engine 16. For example, the exhaust air return duct 172 is configured to return the engine bleed air stream 160 to the turbine engine 16, e.g., the LP compressor 22. The returned engine bleed air stream 160 can be injected back into the turbine engine 16 at any applicable stage of the turbofan engine 10 (e.g., upstream of where the exhaust air supply duct 164 receives the engine bleed air stream 160), or dumped into the bypass airflow passage 56, as described herein.

[0070] In some example embodiments, such as Figure 5AIn accordance with the exemplary embodiments of the present disclosure, the bleed air supply assembly 162 further includes a cavity 168 of the partial-span inlet guide vane 100 in communication with a portion of the engine bleed air stream 160. Further, the partial-span inlet guide vane 100 for the described embodiment further defines a trailing edge opening 170 in airflow communication with the cavity 168, and thus in airflow communication with a portion of the bleed air supply assembly 162. Thus, with this configuration, a portion of the engine bleed air stream 160 can be provided from the bleed air supply assembly 162 to the cavity 168 of the partial-span inlet guide vane 100, and further through the trailing edge opening 170 of the partial-span inlet guide vane 100 during operation of the turbofan engine 10 to reduce the wake created by the respective partial-span inlet guide vane 100.

[0071] In the exemplary embodiments, the bleed air supply assembly 162 further includes a valve 166 (depicted in dashed lines in Figure 5A ) in communication with the bleed air supply duct 164. The valve 166 is transitionable between an open position, in which the bleed air supply duct 164 receives the engine bleed air stream 160 and provides it to the leading edge 108 of the inlet pre-swirl feature, e.g., the inlet pre-swirl feature is configured as a plurality of partial-span inlet guide vanes 100, and a closed position, in which the bleed air supply duct 164 is not in airflow communication with the leading edge 108 of the inlet pre-swirl feature, e.g., the inlet pre-swirl feature is configured as a plurality of partial-span inlet guide vanes 100.

[0072] Referring now to Figure 5B , a cross-sectional view of the partial-span inlet guide vane 100 according to the exemplary embodiments of the present disclosure is provided, showing the position of the bleed air supply duct 164 and the bleed air return duct 172 within the partial-span inlet guide vane 100. Figure 5B The exemplary partial-span inlet guide vane 100 of Figure 5A may be configured in a similar manner as the exemplary partial-span inlet guide vane 100 of Figure 5B . However, for the embodiment of , the exemplary partial-span inlet guide vane 100 does not include the cavity 168 or the trailing edge opening 170.

[0073] For example, the bleed air supply duct 164 is configured to receive the engine bleed air stream 160 and provide it to the leading edge 108 of the partial-span inlet guide vane 100. As shown, the bleed air return duct 172 is positioned between the bleed air supply duct 164 and the trailing edge 110 of the partial-span inlet guide vane 100. Referring to Figure 5B , the partial-span inlet guide vane 100 includes a leading edge 108, a trailing edge 110, a pressure side 120, and a suction side 122.

[0074] Now refer to Figure 6 , provides a close-up cross-sectional view of the forward end of the fan section 14 and turbine 16 of the turbofan engine 10 , according to an exemplary embodiment of the present disclosure. Figure 6 The exemplary engine 10 may be used in the same manner as described above. Figure 5A The exemplary engine is constructed in a similar manner. Figure 6 In the exemplary embodiment of the present invention, the exhaust air supply assembly 162 further includes an exhaust air intermediate duct 174 that is in airflow communication with the exhaust air supply duct 164 and the exhaust air return duct 172 and is disposed between the exhaust air supply duct 164 and the exhaust air return duct 172. The exhaust air intermediate duct 174 is in airflow communication with the upstream end 176 of the outer nacelle 50 and is configured to receive the engine exhaust airflow 160 and provide it to the upstream end 176 of the outer nacelle 50. In this embodiment, the engine exhaust airflow 160 passing through the exhaust air supply duct 164 is the hottest portion of the engine exhaust airflow 160, the engine exhaust airflow 160 passing through the exhaust air intermediate duct 174 is warm air, and the engine exhaust airflow 160 passing through the exhaust air return duct 172 is cooler air that is returned to the turbine 16 (e.g., the LP compressor 22). It is also contemplated that in other exemplary embodiments, the engine exhaust airflow 160 may be directed first to the upstream end 176 of the outer nacelle 50 and then to the leading edge 108 of the partial span inlet guide vanes 100. It is further contemplated that in other exemplary embodiments, the engine exhaust airflow 160 may be directed simultaneously to the upstream end 176 of the outer nacelle 50 and the leading edge 108 of the partial span inlet guide vanes 100.

[0075] In this manner, the engine exhaust airflow 160 is used to heat the partial span inlet guide vanes 100 and the upstream end 176 of the outer nacelle 50. Thus, the engine exhaust airflow 160 is used as a means for reducing ice accumulation or ice formation at the inlet pre-swirl features, for example, configured as a plurality of partial span inlet guide vanes 100, and at the upstream end 176 of the outer nacelle 50.

[0076] Now refer to Figure 7 , provides a close-up cross-sectional view of the forward end of the fan section 14 and turbine 16 of the turbofan engine 10 , according to an exemplary embodiment of the present disclosure. Figure 7 The exemplary engine 10 may be used in the same manner as described above. Figure 5A In the depicted exemplary embodiment of the present disclosure, the heat source 202 comprises a flow of engine oil or engine oil 210. For example, in another exemplary embodiment, with reference to Figure 7In this way, the hot engine oil flow 210 is used to heat the partial-span inlet guide vanes 100 and operates as a device for reducing ice accumulation or ice formation at the inlet pre-swirl feature, e.g., configured as a plurality of partial-span inlet guide vanes 100.

[0077] In this way, the hot engine oil flow 210 is used to heat the partial-span inlet guide vanes 100 and operates as a device for reducing ice accumulation or ice formation at the inlet pre-swirl feature, e.g., configured as a plurality of partial-span inlet guide vanes 100.

[0078] In the depicted example embodiment, the oil supply assembly 262 further includes an oil return conduit 272 in flow communication with the oil supply conduit 264 and the oil source 220. For example, the oil return conduit 272 is configured to return the engine oil flow 210 to the oil source 220.

[0079] In the depicted example embodiment, the oil supply assembly 262 can further include a valve 266 (depicted in dashed line) in communication with the oil supply conduit 264. The valve 266 can be transitionable between an open position, in which the oil supply conduit 264 receives the engine oil flow 210 and provides it to the leading edge 108 of the inlet pre-swirl feature, e.g., configured as a plurality of partial-span inlet guide vanes 100, and a closed position, in which the oil supply conduit 264 is not in airflow communication with the leading edge 108 of the inlet pre-swirl feature, e.g., configured as a plurality of partial-span inlet guide vanes 100.

[0080] Reference is now made to Figure 8A A close-up cross-sectional view of a fan section 14 and a forward end of a turbine 16 of a turbofan engine 10 according to an example embodiment of the present disclosure is provided. Figure 8A An example engine 10 can be configured in a similar manner as the example engine 10 described above. Figure 2 In the depicted example embodiment, the heat source 202 includes an electric heating element 310 disposed proximate to the leading edge 108 of the inlet pre-swirl feature, e.g., configured as a plurality of partial-span inlet guide vanes 100, e.g., closer to the leading edge 108 than to the trailing edge 110. For example, the heat source 202 is more specifically the electric heating element 310 disposed in the leading edge 108 of the inlet pre-swirl feature.

[0081] In this manner, the electric heating element 310 is used to heat the partial span inlet guide vane 100 and operates as a device for reducing ice accumulation or ice formation at the inlet pre-whirl feature, e.g., the inlet pre-whirl feature is configured as a plurality of partial span inlet guide vanes 100.

[0082] In an example embodiment, the heat source 202 further includes an electrical supply assembly 362, the electrical supply assembly 362 including an electrical supply cable 364 of the turbine fan engine 10 in electrical communication with the electric heating element 310.

[0083] Referring to Figure 8B , a cross-sectional view of the partial span inlet guide vane 100 is provided showing the location of the electric heating element 310 within the partial span inlet guide vane 100. For example, the electric heating element 310 includes a first element 312 proximate the leading edge 108, e.g., the first element 312 is closer to the leading edge 108 than the trailing edge 110 of the inlet pre-whirl feature, e.g., the inlet pre-whirl feature is configured as a plurality of partial span inlet guide vanes 100, and a second element 314 positioned within the inlet pre-whirl feature aft of the first element 312, e.g., the inlet pre-whirl feature is configured as a plurality of partial span inlet guide vanes 100. It is contemplated that one or more electric heating elements can be used to heat the partial span inlet guide vane 100. Referring to Figure 8B , the electric heating element 310 can also include a third element 316, a fourth element 318, and a fifth element 320 each positioned within the inlet pre-whirl feature aft of the first element 312, e.g., the inlet pre-whirl feature is configured as a plurality of partial span inlet guide vanes 100.

[0084] Referring to Figure 9A and 9B , in another example embodiment of the present disclosure, the device 200 for reducing ice accumulation or ice formation includes a vibration assembly 410 in mechanical communication with the inlet pre-whirl feature, e.g., the inlet pre-whirl feature is configured as a plurality of partial span inlet guide vanes 100.

[0085] In an example embodiment, the vibration assembly 410 includes a piezoelectric transducer 420 that vibrates the inlet pre-whirl feature, e.g., the inlet pre-whirl feature is configured as a plurality of partial span inlet guide vanes 100. It is contemplated that the vibration assembly 410 can include any other vibration device, such as an ultrasonic vibration device. It is also contemplated that the vibration assembly 410 can be used with any of the heat source embodiments disclosed in Figures 5A-8B .

[0086] In an example embodiment, the vibration assembly 410 further includes an electrical supply assembly 462 including an electrical supply cable 464 of the turbofan engine 10 in electrical communication with the piezoelectric transducer 420.

[0087] Referring Figure 9B , a cross-sectional view of a partial-span inlet guide vane 100 is provided showing the location of the piezoelectric transducer 420 within the partial-span inlet guide vane 100. For example, the piezoelectric transducer 420 is positioned proximate to a leading edge 108 of the partial-span inlet guide vane 100.

[0088] Further, in another example embodiment, the inlet pre-whirl feature, for example, configured as a plurality of partial-span inlet guide vanes 100, is covered with an anti-icing coating 430. For example, in an example embodiment, the anti-icing coating 430 can include an erosion coating or an erosion layer that resists ice accumulation. In an example embodiment, the erosion coating is polyurethane. In an example embodiment, the anti-icing coating 430 has a shore hardness between shore A 50 and shore D 60. In another example embodiment, the anti-icing coating 430 has a shore hardness of shore A 90.

[0089] It is contemplated that the anti-icing coating 430 can cover the inlet pre-whirl feature, for example, configured as a plurality of partial-span inlet guide vanes 100, as shown in Figure 9B However, it is also contemplated that the anti-icing coating 430 can be applied only to selected portions of the inlet pre-whirl feature, for example, configured as a plurality of partial-span inlet guide vanes 100. It is also contemplated that, in the case where the partial-span inlet guide vane 100 is formed from a polymer composite airfoil having a metal leading edge 108, the anti-icing coating 430 is applied only to the portion of the partial-span inlet guide vane 100 rearward of the metal leading edge 108.

[0090] Further aspects of the present disclosure are provided by the subject matter of the following clauses:

[0091] A turbofan engine comprising: a fan including a plurality of fan blades; a turbine machine operably coupled to the fan for driving the fan, the turbine machine including a compressor section, a combustion section, and a turbine section in serial flow order and together defining a core air flowpath; a nacelle surrounding and at least partially enclosing the fan; an inlet pre-whirl feature upstream of the plurality of fan blades, the inlet pre-whirl feature attached to or integrated into the nacelle; and a device on the inlet pre-whirl feature for reducing ice accumulation or ice formation, the device in communication with the inlet pre-whirl feature.

[0092] The turbofan engine of any of the preceding clauses, wherein the device for reducing ice accumulation or ice formation comprises a heat source in thermal communication with the inlet pre-swirl feature.

[0093] The turbofan engine of any of the preceding clauses, wherein the heat source is in communication with a leading edge of the inlet pre-swirl feature.

[0094] The turbofan engine of any of the preceding clauses, wherein the heat source comprises an engine bleed air stream.

[0095] The turbofan engine of any of the preceding clauses, wherein the heat source comprises an air supply assembly comprising a supply duct in airflow communication with a high pressure air source and a leading edge of the inlet pre-swirl feature, the supply duct configured to receive the engine bleed air stream and provide it to the leading edge of the inlet pre-swirl feature.

[0096] The turbofan engine of any of the preceding clauses, wherein the air supply assembly further comprises a return duct in airflow communication with the supply duct and a portion of the turbine, the return duct configured to return the engine bleed air stream to the turbine.

[0097] The turbofan engine of any of the preceding clauses, wherein the air supply assembly further comprises a valve in communication with the supply duct, the valve transitionable between an open position in which the supply duct receives the engine bleed air stream and provides it to the leading edge of the inlet pre-swirl feature and a closed position in which the supply duct is not in airflow communication with the leading edge of the inlet pre-swirl feature.

[0098] The turbofan engine of any of the preceding clauses, wherein the air supply assembly further comprises an intermediate duct in airflow communication with the supply duct and the return duct and disposed between the supply duct and the return duct, the intermediate duct in airflow communication with an upstream end of the nacelle and configured to receive the engine bleed air stream and provide it to the upstream end of the nacelle.

[0099] The turbofan engine of any of the preceding clauses, wherein the high pressure air source is the compressor section of the turbine.

[0100] The turbofan engine of any of the preceding clauses, wherein the heat source comprises an electric heating element disposed in thermal communication with a leading edge of the inlet pre-swirl feature.

[0101] The turbofan engine of any of the preceding clauses, wherein the electric heating element comprises a first element proximate the leading edge of the inlet prewhirl feature, and a second element positioned aft of the first element within the inlet prewhirl feature.

[0102] The turbofan engine of any of the preceding clauses, wherein the heat source comprises an electric supply assembly comprising an electric supply cable in electrical communication with the electric heating element.

[0103] The turbofan engine of any of the preceding clauses, wherein the heat source comprises engine oil.

[0104] The turbofan engine of any of the preceding clauses, wherein the heat source comprises an oil supply assembly comprising an oil supply conduit in flow communication with an oil source and a leading edge of the inlet prewhirl feature, the oil supply conduit configured to receive the engine oil and provide it to the leading edge of the inlet prewhirl feature.

[0105] The turbofan engine of any of the preceding clauses, wherein the oil supply assembly further comprises an oil return conduit in flow communication with the oil supply conduit and the oil source, the oil return conduit configured to return the engine oil to the oil source.

[0106] The turbofan engine of any of the preceding clauses, wherein the oil supply assembly further comprises a valve in communication with the oil supply conduit, the valve transitionable between an open position in which the oil supply conduit receives the engine oil and provides it to the leading edge of the inlet prewhirl feature, and a closed position in which the oil supply conduit is not in flow communication with the leading edge of the inlet prewhirl feature.

[0107] The turbofan engine of any of the preceding clauses, wherein the means for reducing ice accumulation or ice formation comprises a vibration assembly in mechanical communication with the inlet prewhirl feature.

[0108] The turbofan engine of any of the preceding clauses, wherein the means for reducing ice accumulation or ice formation on the inlet prewhirl feature in communication with the inlet prewhirl feature comprises a piezoelectric transducer that vibrates the inlet prewhirl feature.

[0109] The turbofan engine of any of the preceding clauses, wherein the means for reducing ice accumulation or ice formation on the inlet prewhirl feature in communication with the inlet prewhirl feature comprises an anti-icing coating that covers the inlet prewhirl feature.

[0110] The turbofan engine of any of the preceding paragraphs, wherein the inlet prewhirl feature comprises a partial span inlet guide vane forward of the plurality of fan blades of the fan in an axial direction and aft of an inlet of the nacelle.

[0111] A turbofan engine comprising: a fan comprising a plurality of fan blades; a turbine machine operably coupled to the fan for driving the fan, the turbine machine comprising a compressor section, a combustion section, and a turbine section in serial flow order and together defining a core air flowpath; a nacelle surrounding and at least partially enclosing the fan; an inlet prewhirl feature upstream of the plurality of fan blades, the inlet prewhirl feature attached to the nacelle or integrated into the nacelle; and one or more of: a heat source in thermal communication with the inlet prewhirl feature; a vibration assembly in mechanical communication with the inlet prewhirl feature; a piezoelectric transducer to vibrate the inlet prewhirl feature; and an anti-icing coating covering the inlet prewhirl feature.

[0112] This written description uses examples to disclose the disclosure, including the best mode, and also to enable any person skilled in the art to practice the disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the disclosure is defined by the claims, and can include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they 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 languages of the claims.

[0113] While this disclosure has been described as having exemplary designs, the present disclosure can be further modified within the scope of the disclosure. This application is therefore intended to cover any variations, uses, or adaptations of the disclosure using its general principles. Further, this application is intended to cover such departures from the present disclosure as come within known or customary practice in the art to which the disclosure pertains and fall within the limits of the appended claims.

Claims

1. A turbofan engine, characterized in that: include: a fan, the fan comprising a plurality of fan blades; a turbine operatively coupled to the fan for driving the fan, the turbine comprising a compressor section, a combustion section, and a turbine section in serial flow order and together defining a core air flow path; a nacelle surrounding and at least partially enclosing the fan; an inlet pre-swirl feature located upstream of the plurality of fan blades, the inlet pre-swirl feature being attached to or integrated into the nacelle; and means for reducing ice accumulation or ice formation on the inlet pre-swirl feature, the means for reducing ice accumulation or ice formation being in communication with the inlet pre-swirl feature, wherein the means for reducing ice accumulation or ice formation comprises a heat source in thermal communication with the inlet pre-swirl feature, wherein the heat source comprises an engine exhaust gas stream, The heat source includes an air supply assembly, the air supply assembly including: a supply conduit in airflow communication with a high pressure air source and a leading edge of the inlet pre-swirl feature, the supply conduit being configured to receive the engine exhaust airflow and provide it to the leading edge of the inlet pre-swirl feature; a return conduit in communication with the supply conduit and a portion of the turbine's airflow; and an intermediate duct in airflow communication with and disposed between the supply duct and the return duct, the intermediate duct in airflow communication with an upstream end of the nacelle and configured to receive the engine exhaust airflow and provide it to the upstream end of the nacelle, wherein the means for reducing ice accumulation or ice formation comprises an anti-icing coating on the inlet pre-swirl feature, the inlet pre-swirl feature being formed from a polymer composite airfoil having a metal leading edge, wherein the anti-icing coating is applied only to a portion rearward of the metal leading edge.

2. The turbofan engine according to claim 1, characterized in that: Wherein the heat source is in communication with a leading edge of the inlet pre-swirl feature.

3. The turbofan engine according to claim 1, wherein: The return conduit is configured to return the engine exhaust gas flow to the turbine.

4. The turbofan engine according to claim 3, characterized in that: The air supply assembly further includes a valve in communication with the supply conduit, the valve being switchable between an open position in which the supply conduit receives the engine exhaust airflow and provides it to the leading edge of the inlet pre-swirl feature, and a closed position in which the supply conduit is not in communication with the leading edge airflow of the inlet pre-swirl feature.

5. The turbofan engine according to claim 1, wherein: Wherein the high pressure air source is the compressor section of the turbine.

6. The turbofan engine according to claim 1, characterized in that: Wherein the heat source comprises an electric heating element disposed in thermal communication with a leading edge of the inlet pre-swirl feature.

7. The turbofan engine according to claim 6, characterized in that: The electrical heating element includes a first element adjacent the leading edge of the inlet pre-swirl feature, and a second element positioned behind the first element within the inlet pre-swirl feature.

8. The turbofan engine according to claim 6, characterized in that: Wherein the heat source comprises an electrical supply assembly including an electrical supply cable in electrical communication with the electric heating element.

9. The turbofan engine according to claim 1, characterized in that: The heat source includes engine oil.

10. The turbofan engine according to claim 9, characterized in that: wherein the heat source comprises an oil supply assembly including an oil supply conduit in flow communication with an oil source and a leading edge of the inlet pre-swirl feature, the oil supply conduit being configured to receive the engine oil and provide it to the leading edge of the inlet pre-swirl feature.

11. The turbofan engine according to claim 10, characterized in that: The oil supply assembly further includes an oil return conduit in flow communication with the oil supply conduit and the oil source, the oil return conduit being configured to return the engine oil to the oil source.

12. The turbofan engine according to claim 11, characterized in that: The oil supply assembly further includes a valve in communication with the oil supply conduit, the valve being switchable between an open position in which the oil supply conduit receives the engine oil and provides it to the leading edge of the inlet pre-swirl feature, and a closed position in which the oil supply conduit is not in flow communication with the leading edge of the inlet pre-swirl feature.

13. The turbofan engine according to claim 1, wherein: Wherein the means for reducing ice accumulation or ice formation includes a vibrating assembly in mechanical communication with the inlet pre-swirl feature.

14. The turbofan engine according to claim 1, wherein: wherein the means for reducing ice accumulation or ice formation on the inlet pre-swirl feature in communication with the inlet pre-swirl feature comprises a piezoelectric transducer for vibrating the inlet pre-swirl feature.

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