# Patent

## Title
Ice reduction mechanism for turbofan engine

## Abstract

A turbofan engine is provided. The turbofan engine includes a fan comprising a plurality of fan blades; a turbomachine operably coupled to the fan for driving the fan, the turbomachine 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 pre-swirl feature located upstream of the plurality of fan blades, the inlet pre-swirl feature attached to or integrated into the nacelle; and a means for reducing ice buildup or ice formation on the inlet pre-swirl feature, the means in communication with the inlet pre-swirl feature.

## Background
TECHNICAL FIELD
The present subject matter relates generally to a gas turbine engine, or more particularly to a gas turbine engine configured to reduce ice buildup or ice formation on inlet components of the engine.
BACKGROUND
A turbofan engine generally includes a fan having a plurality of fan blades and a turbomachine arranged in flow communication with one another. Additionally, the turbomachine 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 and burned within the combustion section to provide combustion gases. The combustion gases are routed from the combustion section to the turbine section. The flow of combustion gasses through the turbine section drives the turbine section and is then routed through the exhaust section, e.g., to atmosphere.
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, large portions of ice can damage the fan blades or other downstream components of the engine, and may potentially cause an engine flameout.



BRIEF DESCRIPTION OF THE DRAWINGS
A full and enabling disclosure of the present disclosure, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which:
FIG. 1 is a schematic cross-sectional view of an exemplary gas turbine engine according to an exemplary embodiment of the present subject matter.
FIG. 2 is a close-up, schematic, cross-sectional view of a forward end of the exemplary gas turbine engine of FIG. 1 according to an exemplary embodiment of the present subject matter.
FIG. 3 is a schematic view of an inlet to the exemplary gas turbine engine of FIG. 1, along an axial direction of the gas turbine engine of FIG. 1 according to an exemplary embodiment of the present subject matter.
FIG. 4 it is a schematic view of an inlet to a gas turbine engine in accordance with another exemplary embodiment of the present disclosure.
FIG. 5A is a schematic cross-sectional view of an exemplary gas turbine engine according to an exemplary embodiment of the present subject matter.
FIG. 5B is a cross-sectional view of a part span inlet guide vane of an exemplary gas turbine engine at a first location along a span of the part span inlet guide vane according to an exemplary embodiment of the present subject matter.
FIG. 6 is a schematic cross-sectional view of an exemplary gas turbine engine according to another exemplary embodiment of the present subject matter.
FIG. 7 is a schematic cross-sectional view of an exemplary gas turbine engine according to another exemplary embodiment of the present subject matter.
FIG. 8A is a schematic cross-sectional view of an exemplary gas turbine engine according to another exemplary embodiment of the present subject matter.
FIG. 8B is a cross-sectional view of a part span inlet guide vane of an exemplary gas turbine engine at a first location along a span of the part span inlet guide vane according to another exemplary embodiment of the present subject matter.
FIG. 9A is a schematic cross-sectional view of an exemplary gas turbine engine according to another exemplary embodiment of the present subject matter.
FIG. 9B is a cross-sectional view of a part span inlet guide vane of the exemplary gas turbine engine of FIG. 9A at a first location along a span of the part span inlet guide vane according to another exemplary embodiment of the present subject matter.



Corresponding reference characters indicate corresponding parts throughout the several views. The exemplifications set out herein illustrate exemplary embodiments of the disclosure, and such exemplifications are not to be construed as limiting the scope of the disclosure in any manner.
DETAILED DESCRIPTION
Reference will now be made in detail to present embodiments of the disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar designations in the drawings and description have been used to refer to like or similar parts of the disclosure.
The following description is provided to enable those skilled in the art to make and use the described embodiments contemplated for carrying out the disclosure. Various modifications, equivalents, variations, and alternatives, however, will remain readily apparent to those skilled in the art. Any and all such modifications, variations, equivalents, and alternatives are intended to fall within the scope of the present disclosure.
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. Additionally, unless specifically identified otherwise, all embodiments described herein should be considered exemplary.
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 figures. However, it is to be understood that the disclosure may assume various alternative variations, except where expressly specified to the contrary. It is also to be understood that the specific devices illustrated in the attached drawings, and described in the following specification, are simply exemplary embodiments of the disclosure. Hence, specific dimensions and other physical characteristics related to the embodiments disclosed herein are not to be considered as limiting.
As used herein, the terms “first”, “second”, and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components.
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.
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.
The singular forms “a”, “an”, and “the” include plural references unless the context clearly dictates otherwise.
Additionally, the terms “low,” “high,” or their respective comparative degrees (e.g., lower, higher, where applicable) each refer to relative speeds or pressures within an engine, unless otherwise specified. For example, a “low-pressure turbine” operates at a pressure generally lower than a “high-pressure turbine.” Alternatively, unless otherwise specified, the aforementioned terms may be understood in their superlative degree. For example, a “low-pressure turbine” may refer to the lowest maximum pressure turbine within a turbine section, and a “high-pressure turbine” may refer to the highest maximum pressure turbine within the turbine section. An engine of the present disclosure may also include an intermediate pressure turbine, e.g., an engine having three spools.
Approximating language, as used herein throughout the specification and claims, is applied to modify any quantitative representation that could permissibly vary 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 to be limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of an instrument for measuring the value, or the precision of the methods or machines for constructing or manufacturing the components and/or systems. For example, the approximating language may refer to being within a 1, 2, 4, 10, 15, or 20 percent margin. These approximating margins may apply to a single value, either or both endpoints defining numerical ranges, and/or the margin for ranges between endpoints.
Here and throughout the specification and claims, range limitations are combined and interchanged, such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other.
As used herein, the term “fan pressure ratio” refers to a ratio of an air pressure immediately downstream of the fan blades if a fan during operation of the fan to an air pressure immediately upstream of the fan blades of the fan during operation of the fan.
As used herein, the term “rated speed” with reference to a turbofan engine refers to a maximum rotational speed that the turbofan engine may achieve while operating properly. For example, the turbofan engine may be operating at the rated speed during maximum load operations, such as during takeoff operations.
Also as used herein, the term “fan tip speed” as defined by the plurality of fan blades of the fan refers to a linear speed of an outer tip of a fan blade along a radial direction during operation of the fan.
The present disclosure is generally related to a means for reducing ice buildup or ice formation on inlet components of an engine, e.g., an inlet pre-swirl feature configured as a plurality of part span inlet guide vanes.
In some exemplary embodiments of the present disclosure, the means for reducing ice buildup or ice formation includes a heat source that is in thermal communication with the inlet pre-swirl feature, e.g., configured as a plurality of part span inlet guide vanes. For example, in a first exemplary embodiment, the heat source includes an engine bleed airflow. In another exemplary embodiment, the heat source includes an electrical heating element disposed in a leading edge of the inlet pre-swirl feature. In yet another exemplary embodiment, the heat source includes an engine oil.
In other exemplary embodiment, the means for reducing ice buildup or ice formation includes a vibration assembly in mechanical communication with the inlet pre-swirl feature, e.g., configured as a plurality of part span inlet guide vanes, or an anti-icing coating on the pre-swirl feature.
Inclusion of one or more of these means for reducing ice buildup or ice formation provides an anti-icing or de-icing mechanism that prevents the buildup and shedding of pieces of ice into the engine during, e.g., adverse weather conditions, resulting in safer operation of the gas turbine engine.
Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures, FIG. 1 is a schematic cross-sectional view of a gas turbine engine in accordance with an exemplary embodiment of the present disclosure. More particularly, for the embodiment of FIG. 1, the gas turbine engine is an aeronautical, turbofan jet engine 10, referred to herein as “turbofan engine 10”, configured to be mounted to an aircraft, such as in an under-wing configuration or tail-mounted configuration. As shown in FIG. 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 depicted). In general, the turbofan 10 includes a fan section 14 and a turbomachine 16 disposed downstream from the fan section 14 (the turbomachine 16 sometimes also, or alternatively, referred to as a “core turbine engine”).
The exemplary turbomachine 16 depicted generally includes a substantially tubular outer casing 18 that defines an annular inlet 20. The outer casing 18 encases, 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 a jet 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, combustion section 26, turbine section, and jet exhaust nozzle section 32 are arranged in serial flow order and together define a core air flowpath 37 through the turbomachine 16. It is also contemplated that the present disclosure is compatible with an engine having an intermediate pressure turbine, e.g., an engine having three spools.
Referring still the embodiment of FIG. 1, the fan section 14 includes a variable pitch, single stage fan 38, the turbomachine 16 operably coupled to the fan 38 for driving 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 depicted, the fan blades 40 extend outwardly from disk 42 generally along the radial direction R. Each fan blade 40 is rotatable relative to the disk 42 about a pitch axis P by virtue of the fan blades 40 being operatively coupled to a suitable actuation member 44 configured to collectively vary the pitch of the fan blades 40, e.g., in unison. The fan blades 40, disk 42, and actuation member 44 are together rotatable about the longitudinal centerline 12 by LP shaft 36 across a power gear box 46. The power gear box 46 includes a plurality of gears for stepping down the rotational speed of the LP shaft 36 to a more efficient rotational fan speed. Accordingly, for the embodiment depicted, the turbomachine 16 is operably coupled to the fan 38 through the power gear box 46.
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 number of fan blades 40 for a particular application.
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 depicted defines a relatively high fan tip speed and relatively low fan pressure ratio during operation of the turbofan engine at a rated speed. As used herein, the term “fan pressure ratio” refers to a ratio of an air pressure immediately downstream of the fan blades 40 during operation of the fan 38 to an air pressure immediately upstream of the fan blades 40 during operation of the fan 38. For the embodiment depicted in FIG. 1, the fan 38 of the turbofan engine 10 defines a relatively low fan pressure ratio. For example, the turbofan engine 10 depicted defines a fan pressure ratio less than or equal to about 1.5. For example, in certain exemplary embodiments, the turbofan engine 10 may define a fan pressure ratio less than or equal to about 1.4. The fan pressure ratio may 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.
As used herein, the term “rated speed” with reference to the turbofan engine 10 refers to a maximum rotational speed that the turbofan engine 10 may achieve while operating properly. For example, the turbofan engine 10 may be operating at the rated speed during maximum load operations, such as during takeoff operations.
Also as used herein, the term “fan tip speed” defined by the plurality of fan blades 40 refers to a linear speed of an outer tip of a fan blade 40 along the radial direction R during operation of the fan 38. In exemplary embodiments, the turbofan engine 10 of the present disclosure causes the 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 the 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 the rated speed, the fan tip speed of each of the fan blades 40 may 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 the rated speed, the fan tip speed of each of the fan blades 40 may 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.
Referring still to the exemplary embodiment of FIG. 1, the disk 42 is covered by rotatable front nacelle or hub 48 aerodynamically contoured to promote an 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 embodiment depicted, circumferentially, surrounds the fan 38 and at least a portion of the turbomachine 16.
More specifically, 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 turbomachine 16 so as to define a bypass airflow passage 56 therebetween. Additionally, for the embodiment depicted, the outer nacelle 50 is supported relative to the turbomachine 16 by a plurality of circumferentially spaced outlet guide vanes 55.
During operation of the turbofan engine 10, a volume 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 volume of air 58 passes across the fan blades 40, a first portion of the air 58 as indicated by arrows 62 is directed or routed into the bypass airflow passage 56 and a second portion of the air 58 as indicated by arrow 64 is directed or routed into the core air flowpath 37. The ratio between an amount of airflow through the bypass airflow passage 56 (i.e., the first portion of air indicated by arrows 62) to an amount of airflow through the core air flowpath 37 (i.e., the second portion of air indicated by arrows 64) is known as a bypass ratio.
In exemplary embodiments, the bypass ratio during operation of the turbofan engine 10 (e.g., at a 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 a rated speed) may be less than or equal to about ten (10), such as less than or equal to about nine (9). Additionally, the bypass ratio may be at least about two (2).
In other exemplary embodiments, the bypass ratio may generally be between about 7:1 and about 20:1, such as between about 10:1 and about 18:1. The pressure of the second portion of air indicated by arrows 64 is then increased as it is routed through the high pressure (HP) compressor 24 and into the combustion section 26, where it is mixed with fuel and burned to provide combustion gases 66.
In exemplary embodiments, a 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.
Furthermore, the turbofan engine of the present disclosure also provides pre-swirling flow forward of the fan blade tip as described herein.
Referring still to FIG. 1, the compressed second portion of air indicated by arrows 64 from the compressor section mixes with fuel and is burned within the combustion section to provide combustion gases 66. The combustion gases 66 are routed from the combustion section 26, through the HP turbine 28 where a portion of thermal and/or kinetic energy from the combustion gases 66 is extracted via sequential stages of HP turbine stator vanes 68 that are coupled to the outer casing 18 and HP turbine rotor blades 70 that are coupled to the HP shaft 34, thus causing the HP shaft 34 to rotate, thereby supporting operation of the HP compressor 24. The combustion gases 66 are then routed through the LP turbine 30 where a second portion of thermal and kinetic energy is extracted from the combustion gases 66 via sequential stages of LP turbine stator vanes 72 that are coupled to the outer casing 18 and LP turbine rotor blades 74 that are coupled to the LP shaft 36, thus causing the LP shaft 36 to rotate, thereby supporting operation of the LP compressor 22 and/or rotation of the fan 38.
The combustion gases 66 are subsequently routed through the jet exhaust nozzle section 32 of the turbomachine 16 to provide propulsive thrust. Simultaneously, the pressure of the first portion of air indicated by arrows 62 is substantially increased as the first portion of air 62 is routed through the bypass airflow passage 56 before it is exhausted from a fan nozzle exhaust section 76 of the turbofan 10, also providing propulsive thrust. The HP turbine 28, the LP turbine 30, and the jet exhaust nozzle section 32 at least partially define a hot gas path 78 for routing the combustion gases 66 through the turbomachine 16.
Referring still to FIG. 1, the turbofan engine 10 additionally includes a means for reducing ice buildup or ice formation 200 on an inlet pre-swirl feature, e.g., configured as a plurality of part span inlet guide vanes 100, as described in greater detail herein.
In some exemplary embodiments, it will be appreciated that the exemplary turbofan engine 10 of the present disclosure may be a relatively large power class turbofan engine 10. Accordingly, when operated at the rated speed, the turbofan engine 10 may be configured to generate a relatively large amount of thrust. More specifically, when operated at the rated speed, the turbofan engine 10 may 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, e.g., about 150,000 pounds of thrust. Accordingly, the turbofan engine 10 may be referred to as a relatively large power class gas turbine engine.
Moreover, it should be appreciated that the exemplary turbofan engine 10 depicted in FIG. 1 is by way of example only, and that in other exemplary embodiments, the turbofan engine 10 may have any other suitable configuration. For example, in certain exemplary embodiments, the fan may not be a variable pitch fan, the engine may not include a reduction gearbox (e.g., power gearbox 46) driving the fan, may include any other suitable number or arrangement of shafts, spools, compressors, turbines, etc.
As discussed above, the turbofan engine 10 of the present disclosure also provides pre-swirling flow forward a tip of the fan blade 40. Referring now also to FIG. 2, a close-up, cross-sectional view of the fan section 14 and forward end of the turbomachine 16 of the exemplary turbofan engine 10 of FIG. 1 is provided. In exemplary embodiments, the turbofan engine 10 includes an inlet pre-swirl feature located upstream of the plurality of fan blades 40 of the fan 38 and attached to or integrated into the outer nacelle 50. More specifically, for the embodiment of FIGS. 1 and 2, the inlet pre-swirl feature is configured as a plurality of part span inlet guide vanes 100. The plurality of part span inlet guide vanes 100 are each cantilevered from of the outer nacelle 50 (such as from the inner wall 52 of the outer nacelle 50) at a location forward of the plurality of fan blades 40 of the fan 38 along the axial direction A and aft of the inlet 60 of the outer nacelle 50. More specifically, each of the plurality of part span inlet guide vanes 100 define an outer end 102 along the radial direction R, and are attached to/connected to the outer nacelle 50 at the radially outer end 102 through a suitable connection means (not shown). For example, each of the plurality of part span inlet guide vanes 100 may 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.
Further, for the embodiment depicted, the plurality of part span inlet guide vanes 100 extend generally along the radial direction R from the outer end 102 to an inner end 104 (i.e., an inner end 104 along the radial direction R). Moreover, as will be appreciated, for the embodiment depicted, each of the plurality of part span inlet guide vanes 100 are unconnected with an adjacent part span inlet guide vane 100 at the respective inner ends 104 (i.e., adjacent part span inlet guide vanes 100 do not contact one another at the radially inner ends 104, and do not include any intermediate connection members at the radially inner ends 104, such as a connection ring, strut, etc.). More specifically, for the embodiment depicted, each part span inlet guide vane 100 is completely supported by a connection to the outer nacelle 50 at the respective outer end 102 (and not through any structure extending, e.g., between adjacent part span inlet guide vanes 100 at a location inward of the outer end 102 along the radial direction R). As will be discussed below, such may reduce an amount of turbulence generated by the part span inlet guide vanes 100.
Moreover, is depicted, each of the plurality of part span inlet guide vanes 100 do not extend completely between the outer nacelle 50 and, e.g., the hub 48 of the turbofan engine 10. More specifically, for the embodiment depicted, each of the plurality of inlet guide vane define an inlet guide vane (“IGV”) span 106 along the radial direction R, and further each of the plurality of part span inlet guide vanes 100 further define 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 part span inlet guide vane 100 at the leading edge 108 of the part span inlet guide vane 100. Similarly, it will be appreciated, that the plurality of fan blades 40 of the fan 38 define a fan blade span 112 along the radial direction R. More specifically, each of the plurality of fan blades 40 of the fan 38 also 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 fan blade 40 at the leading edge 114 of the respective fan blade 40.
For the embodiment depicted, 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 exemplary embodiments, the IGV span 106 may 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.
Reference will now also be made to FIG. 3, providing an axial view of the inlet 60 to the turbofan engine 10 of FIGS. 1 and 2. As will be appreciated, for the embodiment depicted, the plurality of part span inlet guide vanes 100 of the turbofan engine 10 includes a relatively large number of part span inlet guide vanes 100. More specifically, for the embodiment depicted, the plurality of part span inlet guide vanes 100 includes between about ten part span inlet guide vanes 100 and about fifty part span inlet guide vanes 100. More specifically, for the embodiment depicted, the plurality of part span inlet guide vanes 100 includes between about twenty part span inlet guide vanes 100 and about forty-five part span inlet guide vanes 100, and more specifically, still, the embodiment depicted includes thirty-two part span inlet guide vanes 100. Additionally, for the embodiment depicted, each of the plurality of part span inlet guide vanes 100 are spaced substantially evenly along the circumferential direction C. More specifically, each of the plurality of part span inlet guide vanes 100 defines a circumferential spacing 118 with an adjacent part span inlet guide vane 100, with the circumferential spacing 118 being substantially equal between each adjacent part span inlet guide vane 100.
Although not depicted, in certain exemplary embodiments, the number of part span inlet guide vanes 100 may be substantially equal to the number of fan blades 40 of the fan 38 of the turbofan engine 10. In other embodiments, however, the number of part span 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.
Further, it should be appreciated, that in other exemplary embodiments, the turbofan engine 10 may include any other suitable number of part span inlet guide vanes 100 and/or circumferential spacing 118 of the part span inlet guide vanes 100. For example, referring now briefly to FIG. 4, an axial view of an inlet 60 to a turbofan engine 10 in accordance with another exemplary embodiment of the present disclosure is provided. For the embodiment of FIG. 4, the turbofan engine 10 includes less than twenty part span inlet guide vanes 100. More specifically, for the embodiment of FIG. 4, the turbofan engine 10 includes at least eight part span inlet guide vanes 100, or more specifically includes exactly eight part span inlet guide vanes 100. Additionally, for the embodiment of FIG. 4, the plurality of part span inlet guide vanes 100 are not substantially evenly spaced along the circumferential direction C. For example, at least certain of the plurality of part span inlet guide vanes 100 define a first circumferential spacing 118A, while other of the plurality of part span inlet guide vanes 100 define a second circumferential spacing 118B. For the embodiment depicted, 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 a mean circumferential spacing between adjacent part span inlet guide vanes 100. The non-uniform circumferential spacing may, e.g., offset structure upstream of the part span inlet guide vanes 100.
Referring back to FIG. 2, it will be appreciated that each of the plurality of part span inlet guide vanes 100 is configured to pre-swirl an airflow 58 provided through the inlet 60 of the nacelle 50, upstream of the plurality of fan blades 40 of the fan 38. As briefly discussed above, pre-swirling the airflow 58 provided through the inlet 60 of the nacelle 50 prior to such airflow 58 reaching the plurality of fan blades 40 of the fan 38 may reduce separation losses and/or shock losses, allowing the fan 38 to operate with the relatively high fan tip speeds described above with less losses in efficiency.
As discussed, the present disclosure provides a means for reducing ice buildup or ice formation 200 on the inlet pre-swirl feature, e.g., configured as a plurality of part span inlet guide vanes 100, in communication with the inlet pre-swirl feature. This provides an anti-icing or de-icing mechanism that prevents the buildup and shedding of pieces of ice into the engine during adverse weather conditions.
Referring now generally to FIGS. 5A through 8B, in exemplary embodiments of the present disclosure, the means for reducing ice buildup or ice formation 200 includes a heat source 202 that is in thermal communication with the inlet pre-swirl feature, e.g., configured as a plurality of part span inlet guide vanes 100.
Referring to FIG. 5A, a close-up, cross-sectional view of a fan section 14 and a forward end of a turbomachine 16 of a turbofan engine 10 in accordance with an exemplary embodiment of the present disclosure is provided. The exemplary engine 10 of FIG. 5A may be configured in a similar manner as the exemplary engine of FIG. 2 described above.
In the exemplary embodiment depicted, the heat source 202 includes an engine bleed airflow 160. For example, in a first exemplary embodiment, referring to FIG. 5A, the heat source 202 includes a bleed air supply assembly 162 including a bleed air supply duct 164 in airflow communication with a high pressure air source 165, e.g., engine bleed airflow 160 from the engine 10, and the leading edge 108 of the inlet pre-swirl feature, e.g., configured as a plurality of part span inlet guide vanes 100. The bleed air supply duct 164 is configured to receive and provide the engine bleed airflow 160 to the leading edge 108 of the inlet pre-swirl feature, e.g., configured as a plurality of part span inlet guide vanes 100. For example, the bleed air supply duct 164 may provide the engine bleed airflow 160 to a location proximate the leading edge 108, e.g., to a location closer to the leading edge 108 than the trailing edge 110.
In exemplary embodiments, the high pressure air source 165 is the compressor section, e.g., the LP compressor 22, of the turbomachine 16. For example, in an exemplary embodiment, hot compressor discharge air, e.g., the engine bleed airflow 160, is routed to the leading edge 108 of the part span inlet guide vane 100 and then a cooler engine bleed airflow 160 is returned back to the turbomachine 16. It is contemplated that the compressor discharge air may be sourced from any stage of the LP compressor 22 or the HP compressor 24.
In this manner, the engine bleed airflow 160 is utilized to heat the part span inlet guide vane 100 and operate as a means for reducing ice buildup or ice formation at the inlet pre-swirl feature, e.g., configured as a plurality of part span inlet guide vanes 100.
In an exemplary embodiment, the bleed air supply assembly 162 further includes a bleed air return duct 172 in airflow communication with the bleed air supply duct 164 and a portion of the turbomachine 16. For example, the bleed air return duct 172 is configured to return the engine bleed airflow 160 back to the turbomachine 16, e.g., the LP compressor 22. The return engine bleed airflow 160 may be injected back to the turbomachine 16 at any applicable stage of the turbofan engine 10 (e.g., upstream from where the bleed air supply duct 164 received the engine bleed airflow 160) or dumped into the bypass airflow passage 56 as described herein.
In some exemplary embodiments, such as the exemplary embodiment of FIG. 5A, the bleed air supply assembly 162 further includes a cavity 168 of the part span inlet guide vane 100 in communication with a portion of the engine bleed airflow 160. Furthermore, the part span inlet guide vane 100 for the embodiment depicted further defines a trailing edge opening 170, which is in airflow communication with the cavity 168, and thus is in airflow communication with a portion of the bleed air supply assembly 162. Accordingly, with such a configuration, a portion of the engine bleed airflow 160 may be provided from the bleed air supply assembly 162 to the cavity 168 of the part span inlet guide vane 100, and further through the trailing edge opening 170 of the part span inlet guide vane 100 during operation of the turbofan engine 10 to reduce a wake formed by the respective part span inlet guide vane 100.
In an exemplary embodiment, the bleed air supply assembly 162 further includes a valve 166 (depicted in phantom in FIG. 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 and provides the engine bleed airflow 160 to the leading edge 108 of the inlet pre-swirl feature, e.g., configured as a plurality of part 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., configured as a plurality of part span inlet guide vanes 100.
Referring now to FIG. 5B, providing a cross-sectional view of a part span inlet guide vane 100 in accordance with an exemplary embodiment of the present disclosure, the locations of the bleed air supply duct 164 and the bleed air return duct 172 within the part span inlet guide vane 100 are shown. The exemplary part span inlet guide vane 100 of FIG. 5B may be configured in a similar manner, as the exemplary part span inlet guide vane 100 of FIG. 5A. However, for the embodiment of FIG. 5B, the exemplary part span inlet guide vane 100 does not include a cavity 168 or a trailing edge opening 170.
For example, the bleed air supply duct 164 is configured to receive and provide the engine bleed airflow 160 to the leading edge 108 of the part 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 part span inlet guide vane 100. Referring to FIG. 5B, the part span inlet guide vane 100 includes the leading edge 108, the trailing edge 110, a pressure side 120, and a suction side 122.
Referring now to FIG. 6, a close-up, cross-sectional view of a fan section 14 and a forward end of a turbomachine 16 of a turbofan engine 10 in accordance with an exemplary embodiment of the present disclosure is provided. The exemplary engine 10 of FIG. 6 may be configured in a similar manner as the exemplary engine of FIG. 5A described above. In the exemplary embodiment of FIG. 6, the bleed air supply assembly 162 further includes a bleed air intermediate duct 174 in airflow communication with and disposed between the bleed air supply duct 164 and the bleed air return duct 172. The bleed air intermediate duct 174 is in airflow communication with an upstream end 176 of the outer nacelle 50 and is configured to receive and provide the engine bleed airflow 160 to the upstream end 176 of the outer nacelle 50. In such an embodiment, the engine bleed airflow 160 through the bleed air supply duct 164 is the hottest portion of the engine bleed airflow 160, the engine bleed airflow 160 through the bleed air intermediate duct 174 is a warm air, and the engine bleed airflow 160 through the bleed air return duct 172 is a cooler air return back to the turbomachine 16, e.g., the LP compressor 22. It is also contemplated that in other exemplary embodiments, the engine bleed airflow 160 could first be directed to the upstream end 176 of the outer nacelle 50 and then to the leading edge 108 of the part span inlet guide vane 100. It is further contemplated that in other exemplary embodiments, the engine bleed airflow 160 could be simultaneously directed to the upstream end 176 of the outer nacelle 50 and the leading edge 108 of the part span inlet guide vane 100.
In this manner, the engine bleed airflow 160 is utilized to heat the part span inlet guide vane 100 and the upstream end 176 of the outer nacelle 50. As such, the engine bleed airflow 160 is utilized as a means for reducing ice buildup or ice formation at the inlet pre-swirl feature, e.g., configured as a plurality of part span inlet guide vanes 100, and at the upstream end 176 of the outer nacelle 50.
Referring now to FIG. 7, a close-up, cross-sectional view of a fan section 14 and a forward end of a turbomachine 16 of a turbofan engine 10 in accordance with an exemplary embodiment of the present disclosure is provided. The exemplary engine 10 of FIG. 7 may be configured in a similar manner as the exemplary engine of FIG. 5A described above. In the exemplary embodiment of the present disclosure depicted, the heat source 202 includes an engine oil flow or engine oil 210. For example, in another exemplary embodiment, referring to FIG. 7, the heat source 202 includes an oil supply assembly 262 including an oil supply duct 264 in flow communication with an oil source 220 of the turbofan engine 10 and the leading edge 108 of the inlet pre-swirl feature, e.g., configured as a plurality of part span inlet guide vanes 100. The oil supply duct 264 is configured to receive and provide the engine oil 210 to the leading edge 108 of the inlet pre-swirl feature, e.g., configured as a plurality of part span inlet guide vanes 100.
In this manner, the hot engine oil flow 210 is utilized to heat the part span inlet guide vane 100 and operate as a means for reducing ice buildup or ice formation at the inlet pre-swirl feature, e.g., configured as a plurality of part span inlet guide vanes 100.
In an exemplary embodiment, the oil supply assembly 262 further includes an oil return duct 272 in flow communication with the oil supply duct 264 and the oil source 220. For example, the oil return duct 272 is configured to return the engine oil 210 back to the oil source 220.
In an exemplary embodiment, the oil supply assembly 262 may further include a valve 266 (depicted in phantom) in communication with the oil supply duct 264. The valve 266 is transitionable between an open position in which the oil supply duct 264 receives and provides the engine oil flow 210 to the leading edge 108 of the inlet pre-swirl feature, e.g., configured as a plurality of part span inlet guide vanes 100, and a closed position in which the oil supply duct 264 is not in airflow communication with the leading edge 108 of the inlet pre-swirl feature, e.g., configured as a plurality of part span inlet guide vanes 100.
Referring now to FIG. 8A, a close-up, cross-sectional view of a fan section 14 and a forward end of a turbomachine 16 of a turbofan engine 10 in accordance with an exemplary embodiment of the present disclosure is provided. The exemplary engine 10 of FIG. 8A may be configured in a similar manner as the exemplary engine of FIG. 2 described above. In the exemplary embodiment depicted, the heat source 202 includes an electrical heating element 310 disposed proximate the leading edge 108 of the inlet pre-swirl feature, e.g., configured as a plurality of part span inlet guide vanes 100 (e.g., closer to the leading edge 108 than the trailing edge 110). For example, the heat source 202 is more specifically the electrical heating element 310 disposed in the leading edge 108 of the inlet pre-swirl feature.
In this manner, the electrical heating element 310 is utilized to heat the part span inlet guide vane 100 and operate as a means for reducing ice buildup or ice formation at the inlet pre-swirl feature, e.g., configured as a plurality of part span inlet guide vanes 100.
In an exemplary embodiment, the heat source 202 further includes an electrical supply assembly 362 including an electrical supply cable 364 of the turbofan engine 10 in electrical communication with the electrical heating element 310.
Referring to FIG. 8B, a cross-sectional view of a part span inlet guide vane 100, the locations of an electrical heating element 310 within the part span inlet guide vane 100 are shown. For example, the electrical heating element 310 includes a first element 312 adjacent 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-swirl feature, e.g., configured as a plurality of part span inlet guide vanes 100, and a second element 314 positioned aft of the first element 312 within the inlet pre-swirl feature, e.g., configured as a plurality of part span inlet guide vanes 100. It is contemplated that one or more electrical heating elements may be used to heat the part span inlet guide vane 100. Referring to FIG. 8B, the electrical heating element 310 may also include a third element 316, a fourth element 318, and a fifth element 320 each positioned aft of the first element 312 within the inlet pre-swirl feature, e.g., configured as a plurality of part span inlet guide vanes 100.
Referring to FIGS. 9A and 9B, in another exemplary embodiment of the present disclosure, the means for reducing ice buildup or ice formation 200 includes a vibration assembly 410 in mechanical communication with the inlet pre-swirl feature, e.g., configured as a plurality of part span inlet guide vanes 100.
In an exemplary embodiment, the vibration assembly 410 includes a piezoelectric transducer 420 that vibrates the inlet pre-swirl feature, e.g., configured as a plurality of part span inlet guide vanes 100. It is contemplated that the vibration assembly 410 may include any other vibration devices such as ultrasonic vibration devices. It is also contemplated that the vibration assembly 410 may be used with any of the heat source embodiments disclosed in FIGS. 5A-8B.
In an exemplary 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.
Referring to FIG. 9B, a cross-sectional view of a part span inlet guide vane 100, the location of the piezoelectric transducer 420 within the part span inlet guide vane 100 is shown. For example, the piezoelectric transducer 420 is positioned adjacent the leading edge 108 of the part span inlet guide vane 100.
Furthermore, in another exemplary embodiment, the inlet pre-swirl feature, e.g., configured as a plurality of part span inlet guide vanes 100, is covered with an anti-ice coating 430. For example, in an exemplary embodiment, the anti-ice coating 430 may include an erosion coating or erosion layer that resists ice accumulation. In an exemplary embodiment, the erosion coating is polyurethane. In an exemplary embodiment, the anti-ice coating 430 has a shore hardness between Shore A50 and Shore D60. In another exemplary embodiment, the anti-ice coating 430 has a shore hardness of Shore A90.
It is contemplated that the anti-ice coating 430 may cover the inlet pre-swirl feature, e.g., configured as a plurality of part span inlet guide vanes 100, as shown in FIG. 9B. However, it is also contemplated that the anti-ice coating 430 may only be applied to selected portions of the inlet pre-swirl feature, e.g., configured as a plurality of part span inlet guide vanes 100. It is also contemplated that with a part span inlet guide vane 100 formed of a polymer composite airfoil with a metal leading edge 108, the anti-ice coating 430 is applied only to portions aft of the metal leading edge 108.
Further aspects of the disclosure are provided by the subject matter of the following clauses:
A turbofan engine comprising: a fan comprising a plurality of fan blades; a turbomachine operably coupled to the fan for driving the fan, the turbomachine 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 pre-swirl feature located upstream of the plurality of fan blades, the inlet pre-swirl feature attached to or integrated into the nacelle; and a means for reducing ice buildup or ice formation on the inlet pre-swirl feature, the means in communication with the inlet pre-swirl feature.
The turbofan engine of any preceding clause, wherein the means for reducing ice buildup or ice formation comprises a heat source in thermal communication with the inlet pre-swirl feature.
The turbofan engine of any preceding clause, wherein the heat source is in communication with a leading edge of the inlet pre-swirl feature.
The turbofan engine of any preceding clause, wherein the heat source comprises an engine bleed airflow.
The turbofan engine of any preceding clause, wherein the heat source includes 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 and provide the engine bleed airflow to the leading edge of the inlet pre-swirl feature.
The turbofan engine of any preceding clause, wherein the air supply assembly further comprises a return duct in airflow communication with the supply duct and a portion of the turbomachine, the return duct configured to return the engine bleed airflow back to the turbomachine.
The turbofan engine of any preceding clause, 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 and provides the engine bleed airflow 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.
The turbofan engine of any preceding clause, wherein the air supply assembly further comprises 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 and provide the engine bleed airflow to the upstream end of the nacelle.
The turbofan engine of any preceding clause, wherein the high pressure air source is the compressor section of the turbomachine.
The turbofan engine of any preceding clause, wherein the heat source comprises an electrical heating element disposed in thermal communication with a leading edge of the inlet pre-swirl feature.
The turbofan engine of any preceding clause, wherein the electrical heating element comprises a first element adjacent the leading edge of the inlet pre-swirl feature and a second element positioned aft of the first element within the inlet pre-swirl feature.
The turbofan engine of any preceding clause, wherein the heat source includes an electrical supply assembly comprising an electrical supply cable in electrical communication with the electrical heating element.
The turbofan engine of any preceding clause, wherein the heat source comprises an engine oil.
The turbofan engine of any preceding clause, wherein the heat source includes an oil supply assembly comprising an oil supply duct in flow communication with an oil source and a leading edge of the inlet pre-swirl feature, the oil supply duct configured to receive and provide the engine oil to the leading edge of the inlet pre-swirl feature.
The turbofan engine of any preceding clause, wherein the oil supply assembly further comprises an oil return duct in flow communication with the oil supply duct and the oil source, the oil return duct configured to return the engine oil back to the oil source.
The turbofan engine of any preceding clause, wherein the oil supply assembly further comprises a valve in communication with the oil supply duct, the valve transitionable between an open position in which the oil supply duct receives and provides the engine oil to the leading edge of the inlet pre-swirl feature, and a closed position in which the oil supply duct is not in flow communication with the leading edge of the inlet pre-swirl feature.
The turbofan engine of any preceding clause, wherein the means for reducing ice buildup or ice formation comprises a vibration assembly in mechanical communication with the inlet pre-swirl feature.
The turbofan engine of any preceding clause, wherein the means for reducing ice buildup or ice formation on the inlet pre-swirl feature in communication with the inlet pre-swirl feature comprises a piezoelectric transducer that vibrates the inlet pre-swirl feature.
The turbofan engine of any preceding clause, wherein the means for reducing ice buildup or ice formation on the inlet pre-swirl feature in communication with the inlet pre-swirl feature comprises an anti-ice coating covering the inlet pre-swirl feature.
The turbofan engine of any preceding clause, wherein the inlet pre-swirl feature comprises a part span inlet guide vane at a location forward of the plurality of fan blades of the fan along an axial direction and aft of an inlet of the nacelle.
A turbofan engine comprising: a fan comprising a plurality of fan blades; a turbomachine operably coupled to the fan for driving the fan, the turbomachine 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 pre-swirl feature located upstream of the plurality of fan blades, the inlet pre-swirl feature attached to or integrated into the nacelle; and one or more of: a heat source in thermal communication with the inlet pre-swirl feature; a vibration assembly in mechanical communication with the inlet pre-swirl feature; a piezoelectric transducer that vibrates the inlet pre-swirl feature; and an anti-ice coating covering the inlet pre-swirl feature.
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 may 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.
While this disclosure has been described as having exemplary designs, the present disclosure can be further modified within the scope of this 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 this disclosure pertains and which fall within the limits of the appended claims.

## Claims

1. A turbofan engine comprising:
a fan comprising a plurality of fan blades;
a turbomachine operably coupled to the fan for driving the fan, the turbomachine 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 pre-swirl feature located upstream of the plurality of fan blades, the inlet pre-swirl feature attached to or integrated into the nacelle; and
a means for reducing ice buildup or ice formation on the inlet pre-swirl feature, the means in communication with the inlet pre-swirl feature, wherein the means for reducing ice buildup or ice formation comprises a heat source in thermal communication with the inlet pre-swirl feature, wherein the heat source comprises an engine bleed airflow,
wherein the heat source includes 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 and provide the engine bleed airflow to the leading edge of the inlet pre-swirl feature;
a return duct in airflow communication with the supply duct and a portion of the turbomachine; 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 and provide the engine bleed airflow to the upstream end of the nacelle.

2. The turbofan engine of claim 1, wherein the heat source is in communication with a leading edge of the inlet pre-swirl feature.

3. The turbofan engine of claim 1, wherein the return duct is configured to return the engine bleed airflow back to the turbomachine.

4. The turbofan engine of claim 3, 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 and provides the engine bleed airflow 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.

5. The turbofan engine of claim 1, wherein the high pressure air source is the compressor section of the turbomachine.

6. The turbofan engine of claim 1, wherein the means for reducing ice buildup or ice formation on the inlet pre-swirl feature in communication with the inlet pre-swirl feature comprises an anti-ice coating covering the inlet pre-swirl feature.