US20230265862A1 - Turbofan engine having angled inlet pre-swirl vanes - Google Patents
Turbofan engine having angled inlet pre-swirl vanes Download PDFInfo
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- US20230265862A1 US20230265862A1 US17/676,454 US202217676454A US2023265862A1 US 20230265862 A1 US20230265862 A1 US 20230265862A1 US 202217676454 A US202217676454 A US 202217676454A US 2023265862 A1 US2023265862 A1 US 2023265862A1
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- 238000011144 upstream manufacturing Methods 0.000 claims abstract description 19
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Images
Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
- F01D25/24—Casings; Casing parts, e.g. diaphragms, casing fastenings
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
- F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/04—Air intakes for gas-turbine plants or jet-propulsion plants
- F02C7/045—Air intakes for gas-turbine plants or jet-propulsion plants having provisions for noise suppression
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
- F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
- F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
- F02K3/06—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with front fan
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/40—Casings; Connections of working fluid
- F04D29/52—Casings; Connections of working fluid for axial pumps
- F04D29/54—Fluid-guiding means, e.g. diffusers
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/40—Casings; Connections of working fluid
- F04D29/52—Casings; Connections of working fluid for axial pumps
- F04D29/54—Fluid-guiding means, e.g. diffusers
- F04D29/541—Specially adapted for elastic fluid pumps
- F04D29/542—Bladed diffusers
- F04D29/544—Blade shapes
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
- F04D—NON-POSITIVE-DISPLACEMENT PUMPS
- F04D29/00—Details, component parts, or accessories
- F04D29/66—Combating cavitation, whirls, noise, vibration or the like; Balancing
- F04D29/661—Combating cavitation, whirls, noise, vibration or the like; Balancing especially adapted for elastic fluid pumps
- F04D29/663—Sound attenuation
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2220/00—Application
- F05D2220/30—Application in turbines
- F05D2220/32—Application in turbines in gas turbines
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2220/00—Application
- F05D2220/30—Application in turbines
- F05D2220/36—Application in turbines specially adapted for the fan of turbofan engines
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/10—Stators
- F05D2240/12—Fluid guiding means, e.g. vanes
- F05D2240/127—Vortex generators, turbulators, or the like, for mixing
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2240/00—Components
- F05D2240/10—Stators
- F05D2240/14—Casings or housings protecting or supporting assemblies within
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2250/00—Geometry
- F05D2250/50—Inlet or outlet
- F05D2250/51—Inlet
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/14—Preswirling
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/96—Preventing, counteracting or reducing vibration or noise
- F05D2260/961—Preventing, counteracting or reducing vibration or noise by mistuning rotor blades or stator vanes with irregular interblade spacing, airfoil shape
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Definitions
- the present subject matter relates generally to a gas turbine engine, or more particularly to a gas turbine engine configured to guide an airflow at an inlet of a nacelle and to direct incoming objects towards an outer portion of the engine and away from a core of the engine.
- 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.
- 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.
- 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. 5 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 another exemplary embodiment of the present subject matter.
- FIG. 6 is a cross-sectional view of a part span inlet guide vane of the exemplary gas turbine engine of FIG. 1 at a first location along a span of the part span inlet guide vane.
- FIG. 7 is a cross-sectional view of the part span inlet guide vane of the exemplary gas turbine engine of FIG. 1 at a second location along the span of the part span inlet guide vane.
- FIG. 8 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 another exemplary embodiment of the present subject matter.
- FIG. 9 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 another exemplary embodiment of the present subject matter.
- FIG. 10 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 another exemplary embodiment of the present subject matter.
- FIG. 11 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 another exemplary embodiment of the present subject matter.
- 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.
- 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.
- 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.
- upstream and downstream refer to the relative direction with respect to fluid flow in a fluid pathway.
- upstream refers to the direction from which the fluid flows
- downstream refers to the direction to which the fluid flows.
- the terms “low,” “high,” or their respective comparative degrees each refer to relative speeds or pressures within an engine, unless otherwise specified.
- a “low-pressure turbine” operates at a pressure generally lower than a “high-pressure turbine.”
- the aforementioned terms may be understood in their superlative degree.
- a “low-pressure turbine” may refer to the lowest maximum pressure turbine within a turbine section
- 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 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.
- 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.
- 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.
- 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.
- rated speed with reference to a turbofan engine refers to a maximum rotational speed that the turbofan engine may achieve while operating properly.
- the turbofan engine may be operating at the rated speed during maximum load operations, such as during takeoff operations.
- 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 an inlet pre-swirl vane configured as a plurality of part span inlet guide vanes for a turbofan engine.
- the plurality of part span inlet guide vanes are angled at a first angle with respect to a radius of an outer nacelle of the turbofan engine.
- the plurality of part span inlet guide vanes define a chord and the chord of the part span inlet guide vane is angled at a second angle with respect to the longitudinal axis of the outer nacelle.
- the plurality of part span inlet guide vanes are configured to direct incoming objects towards an outer portion of the turbofan engine.
- the plurality of part span inlet guide vanes are configured to direct incoming objects away from a core air flowpath of the turbofan engine and towards a bypass airflow passage. This provides a deflection mechanism that facilitates ingestion of an object into an outer portion of the turbofan engine by minimizing the chance that the object travels to the core of the turbofan engine.
- Such objects may include bird strikes, hail, ice, sandstorms, debris, and other foreign objects.
- the plurality of part span inlet guide vanes are also configured to pre-swirl an airflow provided through an inlet of the outer nacelle, upstream of the plurality of fan blades of a fan.
- pre-swirling the airflow provided through the inlet of the outer nacelle prior to such airflow reaching the plurality of fan blades of the fan may reduce separation losses and/or shock losses, allowing the fan to operate with relatively high fan tip speeds with less losses in efficiency.
- the plurality of part span inlet guide vanes are also configured to minimize flutter and maintain laminar air flow over the part span inlet guide vanes and minimize turbulence in the inlet.
- 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 engine 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 turbofan engine 10 includes a fan section 14 and a turbomachine 16 disposed downstream from the fan section 14 (the turbomachine
- 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.
- 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 .
- 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.
- 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.
- 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 .
- the fan 38 of the turbofan engine 10 defines a relatively low fan pressure ratio.
- the turbofan engine 10 depicted defines a fan pressure ratio less than or equal to about 1.5.
- the turbofan engine 10 may define a fan pressure ratio less than or equal to about 1.4. In certain exemplary embodiments, it is contemplated that the turbofan engine 10 may define other fan pressure ratios for a particular application.
- 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.
- rated speed refers to a maximum rotational speed that the turbofan engine 10 may achieve while operating properly.
- the turbofan engine 10 may be operating at the rated speed during maximum load operations, such as during takeoff operations.
- 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 circumferential direction during operation of the fan 38 .
- 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.
- the fan tip speed of each of the plurality of fan blades 40 is greater than or equal to 1,000 feet per second and less than or equal to 2,250 feet per second.
- 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 2,250 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 2,250 feet per second. In certain exemplary embodiments, it is contemplated that during operation of the turbofan engine 10 at the rated speed, the fan tip speed of each of the fan blades 40 may define other ranges for a particular application.
- 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 .
- 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 .
- 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 . The outer nacelle 50 includes an inlet 60 at a leading edge 61 of the outer nacelle 50 .
- a volume of air 58 enters the turbofan engine 10 through the inlet 60 of the outer nacelle 50 and/or fan section 14 .
- 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.
- 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 .
- the turbofan engine 10 of the present disclosure also provides pre-swirling flow forward of a tip of the fan blade 40 as described herein.
- the turbofan engine 10 additionally includes an inlet pre-swirl vane, e.g., configured as a plurality of part span inlet guide vanes 100 , as described in greater detail herein.
- 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.
- 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.
- 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.
- the turbofan engine 10 of the present disclosure also provides pre-swirling flow forward a tip of the fan blade 40 .
- 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.
- the turbofan engine 10 includes an inlet pre-swirl vane located upstream of the plurality of fan blades 40 of the fan 38 and coupled to the outer nacelle 50 .
- the inlet pre-swirl vane can be directly attached to, indirectly attached to, or integrated into the outer nacelle 50 . More specifically, for the embodiment of FIGS.
- the inlet pre-swirl vane 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 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 coupled to the outer nacelle 50 at the radially outer end 102 through a suitable connection means (not shown).
- 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 coupled to the outer nacelle 50 in any other suitable manner at the outer end 102 .
- a nacelle assembly 80 of the present disclosure includes the outer nacelle 50 and the inlet pre-swirl vane, e.g., a plurality of part span inlet guide vanes 100 . 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).
- 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.).
- 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 .
- 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 vanes define an inlet guide vane (“IGV”) span 106 along the radial direction R, which 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 . Each of the plurality of part span inlet guide vanes 100 further define a leading edge 108 and a trailing edge 110 .
- IGV inlet guide vane
- the plurality of fan blades 40 of the fan 38 define a fan blade span 112 along the radial direction R, which 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 .
- Each of the plurality of fan blades 40 of the fan 38 also defines a leading edge 114 and a trailing edge 116 .
- 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 .
- 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 .
- 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 .
- the plurality of part span inlet guide vanes 100 of the turbofan engine 10 includes thirty-two part span inlet guide vanes 100 .
- 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 .
- 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 . Additionally, for the embodiment depicted, each of the plurality of part span inlet guide vanes 100 are spaced substantially evenly (e.g., equidistant) 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 .
- the number of part span inlet guide vanes 100 may be substantially equal to the number of fan blades 40 ( FIG. 1 ) of the fan 38 ( FIG. 1 ) 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 .
- 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 .
- 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.
- 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.
- the plurality of part span inlet guide vanes 100 are not substantially evenly spaced along the circumferential direction C.
- at least certain of the plurality of part span inlet guide vanes 100 define a first circumferential spacing 118 A
- other of the plurality of part span inlet guide vanes 100 define a second circumferential spacing 118 B.
- the first circumferential spacing 118 A is at least about twenty percent greater than the second circumferential spacing 118 B, 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.
- 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 .
- the outer nacelle 50 defines a radius NR.
- the plurality of part span inlet guide vanes 100 are angled (e.g., tilted) at a first angle A 1 with respect to the radius NR of the outer nacelle 50 .
- the plurality of part span inlet guide vanes 100 are configured to direct incoming objects towards an outer radial portion of the turbofan engine 10 ( FIG. 1 ).
- the plurality of part span inlet guide vanes 100 are configured to direct incoming objects away from the core air flowpath 37 ( FIG. 1 ) of the turbofan engine 10 and towards a bypass airflow passage 56 ( FIG. 1 ).
- This provides a deflection mechanism that facilitates ingestion of an object into an outer radial portion of the turbofan engine 10 ( FIG. 1 ) and thereby minimizing the chance that the object travels to the core of the turbofan engine 10 .
- the plurality of part span inlet guide vanes 100 are also configured to pre-swirl an airflow 58 ( FIG. 1 ) provided through the inlet 60 of the outer nacelle 50 , upstream of the plurality of fan blades 40 ( FIG. 1 ) of the fan 38 ( FIG. 1 ).
- 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.
- the plurality of part span inlet guide vanes 100 are also configured to minimize flutter and maintain laminar air flow over the part span inlet guide vanes 100 and minimize turbulence in the inlet 60 .
- the first angle A 1 is between approximately 2 degrees and approximately 45 degrees. In other exemplary embodiments, it is contemplated that the first angle A 1 is between other ranges for a particular application.
- the plurality of part span inlet guide vanes 100 are angled at the first angle A 1 with respect to the radius NR of the outer nacelle 50 in a clockwise direction CW, i.e., the same direction that the fan blades 40 ( FIG. 2 ) rotate in.
- the plurality of part span inlet guide vanes 100 are each angled at the same first angle A 1 with respect to the radius NR of the outer nacelle 50 . In other exemplary embodiments, the plurality of part span inlet guide vanes 100 are angled at different first angles A 1 with respect to the radius NR of the outer nacelle 50 as will be described in more detail below.
- FIG. 5 an axial view of the inlet 60 to the turbofan engine 10 of FIGS. 1 and 2 , according to another exemplary embodiment, is provided.
- the plurality of part span inlet guide vanes 100 are angled at the first angle A 1 with respect to the radius NR of the outer nacelle 50 in a counterclockwise direction CCW, i.e., the opposite direction that the fan blades 40 ( FIG. 2 ) rotate in.
- each of the plurality of part span inlet guide vanes 100 are 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 .
- 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 of in efficiency.
- FIG. 6 a cross-sectional view of one part span inlet guide vane 100 along the span of the part span inlet guide vanes 100 , as indicated by Line 6 - 6 in FIG. 2 , is provided.
- the part span inlet guide vane 100 is configured generally as an airfoil having a pressure side 120 and an opposite suction side 122 , and extending between the leading edge 108 and the trailing edge 110 along a camber line 124 .
- the part span inlet guide vane 100 defines a chord line 126 extending directly from the leading edge 108 to the trailing edge 110 .
- chord line 126 of the part span inlet guide vane 100 is angled (e.g., twisted) at a second angle or angle of attack 128 with respect to the longitudinal axis 12 of the outer nacelle 50 ( FIG. 2 ).
- the chord line 126 defines a second angle or angle of attack 128 with an airflow direction 129 of the airflow 58 through the inlet 60 ( FIG. 2 ) of the nacelle 50 .
- the airflow direction 129 is substantially parallel to the axial direction A and the longitudinal axis 12 of the outer nacelle 50 of the turbofan engine 10 .
- the angle of attack 128 at the location depicted along the span 106 of the part span inlet guide vanes 100 is at least approximately five degrees and up to approximately thirty-five degrees.
- the angle of attack 128 at the location depicted along the span 106 of the part span inlet guide vane 100 may be between about ten degrees and about thirty degrees, such as between about fifteen degrees and about twenty-five degrees.
- the part span inlet guide vane 100 at the location depicted along the span 106 ( FIG. 2 ) of the part span inlet guide vane 100 defines a local swirl angle 130 at the trailing edge 110 .
- the “swirl angle” at the trailing edge 110 of the part span inlet guide vane 100 refers to an angle between the airflow direction 129 of the airflow 58 through the inlet 60 ( FIG. 2 ) of the nacelle 50 ( FIG. 2 ) and a reference line 132 defined by a trailing edge section of the pressure side 120 of the part span inlet guide vane 100 .
- the reference line 132 is defined by the aft twenty percent of the pressure side 120 , as measured along the chord line 126 .
- the reference line 132 may be a straight-line average fit of such curve (e.g., using least mean squares).
- a maximum swirl angle 130 refers to the highest swirl angle 130 along the span 106 ( FIG. 2 ) of the part span inlet guide vane 100 .
- the maximum swirl angle 130 is defined proximate the radially outer end 102 ( FIG. 2 ) of the part span inlet guide vane 100 (e.g., at the outer ten percent of the span 106 of the part span inlet guide vanes 100 ), as is represented by the cross-section depicted in FIG. 6 .
- the maximum swirl angle 130 of each part span inlet guide vane 100 at the trailing edge 110 is between approximately five degrees and approximately thirty-five degrees.
- the maximum swirl angle 130 of each part span inlet guide vane 100 at the trailing edge 110 may be between twelve degrees and twenty-five degrees.
- the local swirl angle 130 increases from the radially inner end 104 ( FIG. 2 ) to the radially outer end 102 ( FIG. 2 ) of each part span inlet guide vane 100 .
- FIG. 7 a cross-sectional view of a part span inlet guide vane 100 at a location radially inward from the cross-section viewed in FIG. 6 , as indicated by Line 7 - 7 in FIG. 2 , is provided. As is depicted in FIG.
- the part span inlet guide vane 100 defines the pressure side 120 , the suction side 122 , the leading edge 108 , the trailing edge 110 , the camber line 124 , and chord line 126 .
- the second angle or angle of attack 128 defined by the chord line 126 and the airflow direction 129 of the airflow 58 through the inlet 60 of the nacelle 50 at the location along the span 106 depicted in FIG. 7 is less than the angle of attack 128 at the location along the span 106 depicted in FIG. 6 (e.g., may be at least about twenty percent less, such as at least about fifty percent less, such as up to about one hundred percent less).
- the part span inlet guide vane 100 defines a local swirl angle 130 at the trailing edge 110 at the location along the span 106 of the part span inlet guide vane 100 proximate the inner end 104 , as depicted in FIG. 7 .
- the local swirl angle 130 increases from the radially inner end 104 to the radially outer end 102 of each part span inlet guide vanes 100 .
- the local swirl angle 130 proximate the outer end 102 is greater than the local swirl angle 130 proximate the radially inner end 104 (see FIG. 7 ; e.g., the radially inner ten percent of the span 106 ).
- the local swirl angle 130 may approach zero degrees (e.g., may be less than about five degrees, such as less than about two degrees) at the radially inner end 104 .
- part span inlet guide vanes 100 of such a configuration may reduce an amount of turbulence at the radially inner end 104 ( FIG. 2 ) of each respective part span inlet guide vane 100 .
- such a configuration may provide a desired amount of pre-swirl at the radially outer ends of the plurality of fan blades 40 ( FIG. 2 ) of the fan 38 ( FIG. 2 ) (where the speed of the fan blades 40 is the greatest) to provide a desired reduction in flow separation and/or shock losses that may otherwise occur due to a relatively high speed of the plurality of fan blades 40 at the fan tips during operation of the turbofan engine 10 ( FIG. 2 ).
- FIG. 8 an axial view of the inlet 60 to the turbofan engine 10 of FIGS. 1 and 2 , according to another exemplary embodiment, is provided.
- the plurality of part span inlet guide vanes 100 are angled at different first angles A 1 with respect to the radius NR of the outer nacelle 50 .
- the outer nacelle 50 includes a top portion 210 , a bottom portion 212 , a first side portion 214 , and a second side portion 216 .
- a first portion, e.g., the top portion 210 , of the part span inlet guide vanes 100 are angled at a first angle A 1 with respect to the radius NR of the outer nacelle 50 .
- the part span inlet guide vanes 100 are angled at a first angle A 1 of five degrees or seven degrees with respect to the radius NR of the outer nacelle 50 at the top portion 210 .
- a second portion, e.g., the first side portion 214 , of the part span inlet guide vanes 100 are angled at a different angle with respect to the radius NR of the outer nacelle 50 .
- the part span inlet guide vanes 100 are angled at a first angle A 1 of ten degrees or fifteen degrees with respect to the radius NR of the outer nacelle 50 at the first side portion 214 .
- the plurality of part span inlet guide vanes 100 are angled at different first angles A 1 with respect to the radius NR of the outer nacelle 50 along the circumferential direction C.
- circumferential variation in the angle of tilt can address many issues including cross winds, high angle of attack maneuvers such as takeoff, and engine installation on one side of an aircraft or another.
- cross winds are more likely to affect the 3 or 9 o'clock positions on the engine, e.g., the second side portion 216 and the first side portion 214 , which may make a different angle of tilt more desirable at the 3/9 o'clock positions than at the 6/12 o'clock positions on the engine, e.g., the bottom portion 212 and the top portion 210 .
- first angles A 1 with respect to the radius NR of the outer nacelle 50 may vary for a particular application and may be different at the top portion 210 , the bottom portion 212 , the first side portion 214 , and/or the second side portion 216 of the outer nacelle 50 .
- FIG. 9 an axial view of the inlet 60 to the turbofan engine 10 of FIGS. 1 and 2 , according to another exemplary embodiment, is provided.
- the plurality of part span inlet guide vanes 100 are angled at different first angles A 1 with respect to the radius NR of the outer nacelle 50 .
- the outer nacelle 50 includes a top portion 210 , a bottom portion 212 , a first side portion 214 , and a second side portion 216 .
- the part span inlet guide vanes 100 are angled at different first angles A 1 with respect to the radius NR of the outer nacelle 50 at the top portion 210 and the bottom portion 212 .
- the top portion 210 of the part span inlet guide vanes 100 are angled at a first angle A 1 with respect to the radius NR of the outer nacelle 50 .
- the part span inlet guide vanes 100 are angled at a first angle A 1 of five degrees or seven degrees with respect to the radius NR of the outer nacelle 50 at the top portion 210 .
- the bottom portion 212 of the part span inlet guide vanes 100 are angled at a different angle with respect to the radius NR of the outer nacelle 50 .
- the part span inlet guide vanes 100 are angled at a first angle A 1 of ten degrees with respect to the radius NR of the outer nacelle 50 at the bottom portion 212 .
- the plurality of part span inlet guide vanes 100 are angled at different first angles A 1 with respect to the radius NR of the outer nacelle 50 at the top portion 210 and the bottom portion 212 .
- circumferential variation in the angle of tilt can address many issues including cross winds, high angle of attack maneuvers such as takeoff, and engine installation on one side of an aircraft or another.
- cross winds are more likely to affect the 3 or 9 o'clock positions on the engine, e.g., the second side portion 216 and the first side portion 214 , which may make a different angle of tilt more desirable at the 3/9 o'clock positions than at the 6/12 o'clock positions on the engine, e.g., the bottom portion 212 and the top portion 210 .
- first angles A 1 with respect to the radius NR of the outer nacelle 50 may vary for a particular application and may be different at the top portion 210 , the bottom portion 212 , the first side portion 214 , and/or the second side portion 216 of the outer nacelle 50 .
- FIG. 10 an axial view of the inlet 60 to the turbofan engine 10 of FIGS. 1 and 2 , according to another exemplary embodiment, is provided.
- one of the part span inlet guide vanes 100 are angled at different angles with respect to the radius NR of the outer nacelle 50 at different portions of the part span inlet guide vane 100 .
- the part span inlet guide vane 100 is angled at a first angle A 1 with respect to the radius NR of the outer nacelle 50 a first location 250 of the part span inlet guide vane 100 and the part span inlet guide vane 100 is angled at a second angle A 2 with respect to the radius NR of the outer nacelle 50 at a second location 252 of the part span inlet guide vane 100 .
- the first angle A 1 is different than the second angle A 2 .
- the first angle A 1 is greater than the second angle A 2 .
- varying the angle of tilt e.g., the first angle A 1 and the second angle A 2
- the angle of tilt and sweep may also be used to improve the matching of the swirl that is imparted to the incoming air with the varying linear speed of the fan blade 40 ( FIG.
- the linear speed of the fan blade 40 ( FIG. 2 ) is greatest at the tip, and slowest at the root of the fan blade 40 ( FIG. 2 ), even though the rotational velocity is equal.
- FIG. 11 an axial view of the inlet 60 to the turbofan engine 10 of FIGS. 1 and 2 , according to another exemplary embodiment, is provided.
- one of the part span inlet guide vanes 100 are angled at different angles with respect to the radius NR of the outer nacelle 50 at different locations of the part span inlet guide vane 100 .
- the part span inlet guide vane 100 is angled at a first angle A 1 with respect to the radius NR of the outer nacelle 50 a first location 250 of the part span inlet guide vane 100 and the part span inlet guide vane 100 is angled at a second angle A 2 with respect to the radius NR of the outer nacelle 50 at a second location 252 of the part span inlet guide vane 100 .
- the first angle A 1 is different than the second angle A 2 .
- the first angle A 1 is less than the second angle A 2 .
- varying the angle of tilt e.g., the first angle A 1 and the second angle A 2
- the angle of tilt and sweep may also be used to improve the matching of the swirl that is imparted to the incoming air with the varying linear speed of the fan blade 40 ( FIG.
- the linear speed of the fan blade 40 ( FIG. 2 ) is greatest at the tip, and slowest at the root of the fan blade 40 ( FIG. 2 ), even though the rotational velocity is equal.
- 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, the nacelle defining a radius and a longitudinal axis; and an inlet pre-swirl vane located upstream of the plurality of fan blades and defining a chord, the inlet pre-swirl vane coupled to the nacelle, wherein the inlet pre-swirl vane is angled at a first angle with respect to the radius of the nacelle, and wherein the chord of the inlet pre-swirl vane is angled at a second angle with respect to the longitudinal axis of the nacelle.
- turbofan engine of any preceding clause wherein the inlet pre-swirl vane is angled at the first angle with respect to the radius of the nacelle in a clockwise direction from an inlet of the nacelle.
- turbofan engine of any preceding clause wherein the inlet pre-swirl vane is angled at the first angle with respect to the radius of the nacelle in a counterclockwise direction from an inlet of the nacelle.
- turbofan engine of any preceding clause wherein the first angle is between approximately 2 degrees and approximately 45 degrees.
- turbofan engine of any preceding clause wherein the second angle is between approximately 5 degrees and approximately 35 degrees.
- the inlet pre-swirl vane is one of a plurality of part span inlet guide vanes extending from the nacelle upstream of the plurality of fan blades and aft of an inlet of the nacelle.
- each of the plurality of part span inlet guide vanes is angled at the same first angle with respect to the radius of the nacelle.
- turbofan engine of any preceding clause wherein a first portion of each of the plurality of part span inlet guide vanes is angled at the first angle with respect to the radius of the nacelle, wherein a second portion of each of the plurality of part span inlet guide vanes is angled at a third angle with respect to the radius of the nacelle, and wherein the first angle is different than the third angle.
- the nacelle includes a top portion, a bottom portion, a first side portion, and a second side portion, wherein a first portion of each of the plurality of part span inlet guide vanes is angled at the first angle with respect to the radius of the nacelle at the top portion, wherein a second portion of each of the plurality of part span inlet guide vanes is angled at a third angle with respect to the radius of the nacelle at the bottom portion, and wherein the first angle is different than the third angle.
- a nacelle assembly for a turbofan engine comprising a fan including a plurality of fan blades
- the nacelle assembly comprising: a nacelle surrounding and at least partially enclosing the fan, the nacelle defining a radius and a longitudinal axis; and an inlet pre-swirl vane located upstream of the plurality of fan blades and defining a chord, the inlet pre-swirl vane coupled to the nacelle, wherein the inlet pre-swirl vane is angled at a first angle with respect to the radius of the nacelle, and wherein the chord of the inlet pre-swirl vane is angled at a second angle with respect to the longitudinal axis of the nacelle.
- the inlet pre-swirl vane is one of a plurality of part span inlet guide vanes extending from the nacelle upstream of the plurality of fan blades and aft of an inlet of the nacelle.
- each of the plurality of part span inlet guide vanes is angled at the same first angle with respect to the radius of the nacelle.
- a first portion of each of the plurality of part span inlet guide vanes is angled at the first angle with respect to the radius of the nacelle, wherein a second portion of each of the plurality of part span inlet guide vanes is angled at a third angle with respect to the radius of the nacelle, and wherein the first angle is different than the third angle.
- the nacelle includes a top portion, a bottom portion, a first side portion, and a second side portion, wherein a first portion of each of the plurality of part span inlet guide vanes is angled at the first angle with respect to the radius of the nacelle at the top portion, wherein a second portion of each of the plurality of part span inlet guide vanes is angled at a third angle with respect to the radius of the nacelle at the bottom portion, and wherein the first angle is different than the third angle.
- a nacelle assembly for a turbofan engine comprising a fan including a plurality of fan blades, the nacelle assembly comprising: a nacelle surrounding and at least partially enclosing the fan, the nacelle defining a radius; and an inlet pre-swirl vane located upstream of the plurality of fan blades, the inlet pre-swirl vane coupled to the nacelle, wherein the inlet pre-swirl vane is angled at a first angle with respect to the radius of the nacelle at a first location of the inlet pre-swirl vane, wherein the inlet pre-swirl vane is angled at a second angle with respect to the radius of the nacelle at a second location of the inlet pre-swirl vane, and wherein the first angle is different than the second angle.
- nacelle assembly of any preceding clause, wherein the nacelle defines a longitudinal axis, wherein the inlet pre-swirl vane defines a chord, and wherein the chord of the inlet pre-swirl vane is angled at a third angle with respect to the longitudinal axis of the nacelle.
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Abstract
Description
- The present subject matter relates generally to a gas turbine engine, or more particularly to a gas turbine engine configured to guide an airflow at an inlet of a nacelle and to direct incoming objects towards an outer portion of the engine and away from a core of the engine.
- 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.
- 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 ofFIG. 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 ofFIG. 1 , along an axial direction of the gas turbine engine ofFIG. 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. 5 is a schematic view of an inlet to the exemplary gas turbine engine ofFIG. 1 , along an axial direction of the gas turbine engine ofFIG. 1 according to another exemplary embodiment of the present subject matter. -
FIG. 6 is a cross-sectional view of a part span inlet guide vane of the exemplary gas turbine engine ofFIG. 1 at a first location along a span of the part span inlet guide vane. -
FIG. 7 is a cross-sectional view of the part span inlet guide vane of the exemplary gas turbine engine ofFIG. 1 at a second location along the span of the part span inlet guide vane. -
FIG. 8 is a schematic view of an inlet to the exemplary gas turbine engine ofFIG. 1 , along an axial direction of the gas turbine engine ofFIG. 1 according to another exemplary embodiment of the present subject matter. -
FIG. 9 is a schematic view of an inlet to the exemplary gas turbine engine ofFIG. 1 , along an axial direction of the gas turbine engine ofFIG. 1 according to another exemplary embodiment of the present subject matter. -
FIG. 10 is a schematic view of an inlet to the exemplary gas turbine engine ofFIG. 1 , along an axial direction of the gas turbine engine ofFIG. 1 according to another exemplary embodiment of the present subject matter. -
FIG. 11 is a schematic view of an inlet to the exemplary gas turbine engine ofFIG. 1 , along an axial direction of the gas turbine engine ofFIG. 1 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.
- 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 an inlet pre-swirl vane configured as a plurality of part span inlet guide vanes for a turbofan engine. In the present disclosure, the plurality of part span inlet guide vanes are angled at a first angle with respect to a radius of an outer nacelle of the turbofan engine. Furthermore, the plurality of part span inlet guide vanes define a chord and the chord of the part span inlet guide vane is angled at a second angle with respect to the longitudinal axis of the outer nacelle.
- In this manner, the plurality of part span inlet guide vanes are configured to direct incoming objects towards an outer portion of the turbofan engine. The plurality of part span inlet guide vanes are configured to direct incoming objects away from a core air flowpath of the turbofan engine and towards a bypass airflow passage. This provides a deflection mechanism that facilitates ingestion of an object into an outer portion of the turbofan engine by minimizing the chance that the object travels to the core of the turbofan engine. Such objects may include bird strikes, hail, ice, sandstorms, debris, and other foreign objects.
- Furthermore, in this manner, the plurality of part span inlet guide vanes are also configured to pre-swirl an airflow provided through an inlet of the outer nacelle, upstream of the plurality of fan blades of a fan. As discussed herein, pre-swirling the airflow provided through the inlet of the outer nacelle prior to such airflow reaching the plurality of fan blades of the fan may reduce separation losses and/or shock losses, allowing the fan to operate with relatively high fan tip speeds with less losses in efficiency. Furthermore, in this manner, the plurality of part span inlet guide vanes are also configured to minimize flutter and maintain laminar air flow over the part span inlet guide vanes and minimize turbulence in the inlet.
- 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 ofFIG. 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 inFIG. 1 , theturbofan engine 10 defines an axial direction A (extending parallel to alongitudinal 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, theturbofan engine 10 includes afan section 14 and aturbomachine 16 disposed downstream from the fan section 14 (theturbomachine 16 sometimes also, or alternatively, referred to as a “core turbine engine”). - The
exemplary turbomachine 16 depicted generally includes a substantially tubularouter casing 18 that defines anannular inlet 20. Theouter 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; acombustion section 26; a turbine section including a first, high pressure (HP)turbine 28 and a second, low pressure (LP) turbine 30; and a jetexhaust nozzle section 32. A high pressure (HP)shaft 34 drivingly connects theHP turbine 28 to theHP compressor 24. A low pressure (LP)shaft 36 drivingly connects the LP turbine 30 to theLP compressor 22. The compressor section,combustion section 26, turbine section, and jetexhaust nozzle section 32 are arranged in serial flow order and together define acore air flowpath 37 through theturbomachine 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 to the embodiment of
FIG. 1 , thefan section 14 includes a variable pitch,single stage fan 38, theturbomachine 16 operably coupled to thefan 38 for driving thefan 38. Thefan 38 includes a plurality ofrotatable fan blades 40 coupled to adisk 42 in a spaced apart manner. As depicted, thefan blades 40 extend outwardly fromdisk 42 generally along the radial direction R. Eachfan blade 40 is rotatable relative to thedisk 42 about a pitch axis P by virtue of thefan blades 40 being operatively coupled to asuitable actuation member 44 configured to collectively vary the pitch of thefan blades 40, e.g., in unison. Thefan blades 40,disk 42, andactuation member 44 are together rotatable about thelongitudinal centerline 12 byLP shaft 36 across apower gear box 46. Thepower gear box 46 includes a plurality of gears for stepping down the rotational speed of theLP shaft 36 to a more efficient rotational fan speed. Accordingly, for the embodiment depicted, theturbomachine 16 is operably coupled to thefan 38 through thepower gear box 46. - In exemplary embodiments, the
fan section 14 includes twenty-two (22) orfewer fan blades 40. In certain exemplary embodiments, thefan section 14 includes twenty (20) orfewer fan blades 40. In certain exemplary embodiments, thefan section 14 includes eighteen (18) orfewer fan blades 40. In certain exemplary embodiments, thefan section 14 includes sixteen (16) orfewer fan blades 40. In certain exemplary embodiments, it is contemplated that thefan section 14 includes other number offan blades 40 for a particular application. - During operation of the
turbofan engine 10, thefan 38 defines a fan pressure ratio and the plurality offan blades 40 each define a fan tip speed. Theexemplary 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 thefan blades 40 during operation of thefan 38 to an air pressure immediately upstream of thefan blades 40 during operation of thefan 38. For the embodiment depicted inFIG. 1 , thefan 38 of theturbofan engine 10 defines a relatively low fan pressure ratio. For example, theturbofan engine 10 depicted defines a fan pressure ratio less than or equal to about 1.5. For example, in certain exemplary embodiments, theturbofan engine 10 may define a fan pressure ratio less than or equal to about 1.4. In certain exemplary embodiments, it is contemplated that theturbofan engine 10 may define other fan pressure ratios for a particular application. The fan pressure ratio may be the fan pressure ratio of thefan 38 during operation of theturbofan engine 10, such as during operation of theturbofan 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 theturbofan engine 10 may achieve while operating properly. For example, theturbofan 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 afan blade 40 along the circumferential direction during operation of thefan 38. In exemplary embodiments, theturbofan engine 10 of the present disclosure causes thefan blades 40 of thefan 38 to rotate at a relatively high rotational speed. For example, during operation of theturbofan engine 10 at the rated speed, the fan tip speed of each of the plurality offan blades 40 is greater than or equal to 1,000 feet per second and less than or equal to 2,250 feet per second. In certain exemplary embodiments, during operation of theturbofan engine 10 at the rated speed, the fan tip speed of each of thefan blades 40 may be greater than or equal to 1,250 feet per second and less than or equal to 2,250 feet per second. In certain exemplary embodiments, during operation of theturbofan engine 10 at the rated speed, the fan tip speed of each of thefan 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 2,250 feet per second. In certain exemplary embodiments, it is contemplated that during operation of theturbofan engine 10 at the rated speed, the fan tip speed of each of thefan blades 40 may define other ranges for a particular application. - Referring still to the exemplary embodiment of
FIG. 1 , thedisk 42 is covered by rotatable front nacelle orhub 48 aerodynamically contoured to promote an airflow through the plurality offan blades 40. Additionally, theexemplary fan section 14 includes an annular fan casing orouter nacelle 50 that at least partially, and for the embodiment depicted, circumferentially, surrounds thefan 38 and at least a portion of theturbomachine 16. - More specifically, the
outer nacelle 50 includes aninner wall 52 and adownstream section 54 of theinner wall 52 of theouter nacelle 50 extends over an outer portion of theturbomachine 16 so as to define abypass airflow passage 56 therebetween. Additionally, for the embodiment depicted, theouter nacelle 50 is supported relative to theturbomachine 16 by a plurality of circumferentially spaced outlet guide vanes 55. Theouter nacelle 50 includes aninlet 60 at aleading edge 61 of theouter nacelle 50. - During operation of the
turbofan engine 10, a volume ofair 58 enters theturbofan engine 10 through theinlet 60 of theouter nacelle 50 and/orfan section 14. As the volume ofair 58 passes across thefan blades 40, a first portion of theair 58 as indicated byarrows 62 is directed or routed into thebypass airflow passage 56 and a second portion of theair 58 as indicated byarrow 64 is directed or routed into thecore 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. - Referring still to
FIG. 1 , the compressed second portion of air indicated byarrows 64 from the compressor section mixes with fuel and is burned within the combustion section to providecombustion gases 66. Thecombustion gases 66 are routed from thecombustion section 26, through theHP turbine 28 where a portion of thermal and/or kinetic energy from thecombustion gases 66 is extracted via sequential stages of HPturbine stator vanes 68 that are coupled to theouter casing 18 and HPturbine rotor blades 70 that are coupled to theHP shaft 34, thus causing theHP shaft 34 to rotate, thereby supporting operation of theHP compressor 24. Thecombustion gases 66 are then routed through the LP turbine 30 where a second portion of thermal and kinetic energy is extracted from thecombustion gases 66 via sequential stages of LPturbine stator vanes 72 that are coupled to theouter casing 18 and LPturbine rotor blades 74 that are coupled to theLP shaft 36, thus causing theLP shaft 36 to rotate, thereby supporting operation of theLP compressor 22 and/or rotation of thefan 38. - The
combustion gases 66 are subsequently routed through the jetexhaust nozzle section 32 of theturbomachine 16 to provide propulsive thrust. Simultaneously, the pressure of the first portion of air indicated byarrows 62 is substantially increased as the first portion ofair 62 is routed through thebypass airflow passage 56 before it is exhausted from a fannozzle exhaust section 76 of theturbofan 10, also providing propulsive thrust. TheHP turbine 28, the LP turbine 30, and the jetexhaust nozzle section 32 at least partially define ahot gas path 78 for routing thecombustion gases 66 through theturbomachine 16. - Referring still to
FIG. 1 , theturbofan engine 10 of the present disclosure also provides pre-swirling flow forward of a tip of thefan blade 40 as described herein. For example, theturbofan engine 10 additionally includes an inlet pre-swirl vane, e.g., configured as a plurality of part spaninlet 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 powerclass turbofan engine 10. Accordingly, when operated at the rated speed, theturbofan engine 10 may be configured to generate a relatively large amount of thrust. More specifically, when operated at the rated speed, theturbofan 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, theturbofan 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 inFIG. 1 is by way of example only, and that in other exemplary embodiments, theturbofan 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 thefan blade 40. Referring now also toFIG. 2 , a close-up, cross-sectional view of thefan section 14 and forward end of theturbomachine 16 of theexemplary turbofan engine 10 ofFIG. 1 is provided. In exemplary embodiments, theturbofan engine 10 includes an inlet pre-swirl vane located upstream of the plurality offan blades 40 of thefan 38 and coupled to theouter nacelle 50. For example, the inlet pre-swirl vane can be directly attached to, indirectly attached to, or integrated into theouter nacelle 50. More specifically, for the embodiment ofFIGS. 1 and 2 , the inlet pre-swirl vane is configured as a plurality of part span inlet guide vanes 100. The plurality of part spaninlet guide vanes 100 are each cantilevered from the outer nacelle 50 (such as from theinner wall 52 of the outer nacelle 50) at a location forward of the plurality offan blades 40 of thefan 38 along the axial direction A and aft of theinlet 60 of theouter nacelle 50. More specifically, each of the plurality of part spaninlet guide vanes 100 define anouter end 102 along the radial direction R, and are coupled to theouter nacelle 50 at the radiallyouter end 102 through a suitable connection means (not shown). For example, each of the plurality of part spaninlet guide vanes 100 may be bolted to theinner wall 52 of theouter nacelle 50 at theouter end 102, welded to theinner wall 52 of theouter nacelle 50 at theouter end 102, or coupled to theouter nacelle 50 in any other suitable manner at theouter end 102. - Referring still to
FIG. 2 , in an exemplary embodiment, anacelle assembly 80 of the present disclosure includes theouter nacelle 50 and the inlet pre-swirl vane, e.g., a plurality of part span inlet guide vanes 100. Further, for the embodiment depicted, the plurality of part spaninlet guide vanes 100 extend generally along the radial direction R from theouter end 102 to an inner end 104 (i.e., aninner end 104 along the radial direction R). Moreover, as will be appreciated, for the embodiment depicted, each of the plurality of part spaninlet guide vanes 100 are unconnected with an adjacent part spaninlet guide vane 100 at the respective inner ends 104 (i.e., adjacent part spaninlet 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 spaninlet guide vane 100 is completely supported by a connection to theouter nacelle 50 at the respective outer end 102 (and not through any structure extending, e.g., between adjacent part spaninlet guide vanes 100 at a location inward of theouter 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, as depicted, each of the plurality of part span
inlet guide vanes 100 do not extend completely between theouter nacelle 50 and, e.g., thehub 48 of theturbofan engine 10. More specifically, for the embodiment depicted, each of the plurality of inlet guide vanes define an inlet guide vane (“IGV”)span 106 along the radial direction R, which refers to a measure along the radial direction R between theouter end 102 and theinner end 104 of the part spaninlet guide vane 100 at theleading edge 108 of the part spaninlet guide vane 100. Each of the plurality of part spaninlet guide vanes 100 further define aleading edge 108 and a trailingedge 110. Similarly, it will be appreciated, that the plurality offan blades 40 of thefan 38 define afan blade span 112 along the radial direction R, which refers to a measure along the radial direction R between a radially outer tip and a base of thefan blade 40 at theleading edge 114 of therespective fan blade 40. Each of the plurality offan blades 40 of thefan 38 also defines aleading edge 114 and a trailingedge 116. - For the embodiment depicted, the
IGV span 106 is at least about five percent of thefan blade span 112 and up to about fifty-five percent of thefan blade span 112. For example, in certain exemplary embodiments, theIGV span 106 may be between about fifteen percent of thefan blade span 112 and about forty-five percent of thefan blade span 112, such as between about thirty percent of thefan blade span 112 and about forty percent of thefan blade span 112. - Reference will now also be made to
FIG. 3 , providing an axial view of theinlet 60 to theturbofan engine 10 ofFIGS. 1 and 2 . As will be appreciated, the plurality of part spaninlet guide vanes 100 of theturbofan engine 10 includes a relatively large number of part span inlet guide vanes 100. For example, for the embodiment depicted, the plurality of part spaninlet guide vanes 100 of theturbofan engine 10 includes thirty-two part span inlet guide vanes 100. In other exemplary embodiments, it is contemplated that the plurality of part spaninlet guide vanes 100 includes between about ten part spaninlet guide vanes 100 and about fifty part span inlet guide vanes 100. In further exemplary embodiments, it is contemplated that the plurality of part spaninlet guide vanes 100 includes between about twenty part spaninlet guide vanes 100 and about forty-five part span inlet guide vanes 100. Additionally, for the embodiment depicted, each of the plurality of part spaninlet guide vanes 100 are spaced substantially evenly (e.g., equidistant) along the circumferential direction C. More specifically, each of the plurality of part spaninlet guide vanes 100 defines acircumferential spacing 118 with an adjacent part spaninlet guide vane 100, with thecircumferential spacing 118 being substantially equal between each adjacent part spaninlet 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 (FIG. 1 ) of the fan 38 (FIG. 1 ) of theturbofan engine 10. In other embodiments, however, the number of part spaninlet guide vanes 100 may be greater than the number offan blades 40 of thefan 38 of theturbofan engine 10, or alternatively, may be less than the number offan blades 40 of thefan 38 of theturbofan engine 10. - Further, it should be appreciated, that in other exemplary embodiments, the
turbofan engine 10 may include any other suitable number of part spaninlet guide vanes 100 and/orcircumferential spacing 118 of the part span inlet guide vanes 100. For example, referring now briefly toFIG. 4 , an axial view of aninlet 60 to aturbofan engine 10 in accordance with another exemplary embodiment of the present disclosure is provided. For the embodiment ofFIG. 4 , theturbofan engine 10 includes less than twenty part span inlet guide vanes 100. More specifically, for the embodiment ofFIG. 4 , theturbofan engine 10 includes at least eight part spaninlet guide vanes 100, or more specifically includes exactly eight part span inlet guide vanes 100. Additionally, for the embodiment ofFIG. 4 , the plurality of part spaninlet guide vanes 100 are not substantially evenly spaced along the circumferential direction C. For example, at least certain of the plurality of part spaninlet guide vanes 100 define a firstcircumferential spacing 118A, while other of the plurality of part spaninlet guide vanes 100 define a secondcircumferential spacing 118B. For the embodiment depicted, the firstcircumferential spacing 118A is at least about twenty percent greater than the secondcircumferential 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, thecircumferential 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. 3 , theouter nacelle 50 defines a radius NR. For the embodiment depicted, the plurality of part spaninlet guide vanes 100 are angled (e.g., tilted) at a first angle A1 with respect to the radius NR of theouter nacelle 50. - In this manner, the plurality of part span
inlet guide vanes 100 are configured to direct incoming objects towards an outer radial portion of the turbofan engine 10 (FIG. 1 ). The plurality of part spaninlet guide vanes 100 are configured to direct incoming objects away from the core air flowpath 37 (FIG. 1 ) of theturbofan engine 10 and towards a bypass airflow passage 56 (FIG. 1 ). This provides a deflection mechanism that facilitates ingestion of an object into an outer radial portion of the turbofan engine 10 (FIG. 1 ) and thereby minimizing the chance that the object travels to the core of theturbofan engine 10. - Furthermore, in this manner, the plurality of part span
inlet guide vanes 100 are also configured to pre-swirl an airflow 58 (FIG. 1 ) provided through theinlet 60 of theouter nacelle 50, upstream of the plurality of fan blades 40 (FIG. 1 ) of the fan 38 (FIG. 1 ). As discussed herein, pre-swirling theairflow 58 provided through theinlet 60 of thenacelle 50 prior tosuch airflow 58 reaching the plurality offan blades 40 of thefan 38 may reduce separation losses and/or shock losses, allowing thefan 38 to operate with the relatively high fan tip speeds described above with less losses in efficiency. Furthermore, in this manner, the plurality of part spaninlet guide vanes 100 are also configured to minimize flutter and maintain laminar air flow over the part spaninlet guide vanes 100 and minimize turbulence in theinlet 60. - In certain exemplary embodiments, the first angle A1 is between approximately 2 degrees and approximately 45 degrees. In other exemplary embodiments, it is contemplated that the first angle A1 is between other ranges for a particular application.
- In the exemplary embodiment depicted in
FIG. 3 , the plurality of part spaninlet guide vanes 100 are angled at the first angle A1 with respect to the radius NR of theouter nacelle 50 in a clockwise direction CW, i.e., the same direction that the fan blades 40 (FIG. 2 ) rotate in. - In an exemplary embodiment, the plurality of part span
inlet guide vanes 100 are each angled at the same first angle A1 with respect to the radius NR of theouter nacelle 50. In other exemplary embodiments, the plurality of part spaninlet guide vanes 100 are angled at different first angles A1 with respect to the radius NR of theouter nacelle 50 as will be described in more detail below. - Referring now to
FIG. 5 , an axial view of theinlet 60 to theturbofan engine 10 ofFIGS. 1 and 2 , according to another exemplary embodiment, is provided. Referring toFIG. 5 , in another exemplary embodiment, the plurality of part spaninlet guide vanes 100 are angled at the first angle A1 with respect to the radius NR of theouter nacelle 50 in a counterclockwise direction CCW, i.e., the opposite direction that the fan blades 40 (FIG. 2 ) rotate in. - Referring now back to
FIG. 2 , as described above, each of the plurality of part spaninlet guide vanes 100 are configured to pre-swirl anairflow 58 provided through theinlet 60 of thenacelle 50, upstream of the plurality offan blades 40 of thefan 38. As discussed above, pre-swirling theairflow 58 provided through theinlet 60 of thenacelle 50 prior tosuch airflow 58 reaching the plurality offan blades 40 of thefan 38 may reduce separation losses and/or shock losses, allowing thefan 38 to operate with the relatively high fan tip speeds described above with less losses of in efficiency. - For example, referring first to
FIG. 6 , a cross-sectional view of one part spaninlet guide vane 100 along the span of the part spaninlet guide vanes 100, as indicated by Line 6-6 inFIG. 2 , is provided. As is depicted, the part spaninlet guide vane 100 is configured generally as an airfoil having apressure side 120 and anopposite suction side 122, and extending between theleading edge 108 and the trailingedge 110 along acamber line 124. Additionally, the part spaninlet guide vane 100 defines achord line 126 extending directly from theleading edge 108 to the trailingedge 110. Thechord line 126 of the part spaninlet guide vane 100 is angled (e.g., twisted) at a second angle or angle ofattack 128 with respect to thelongitudinal axis 12 of the outer nacelle 50 (FIG. 2 ). For example, thechord line 126 defines a second angle or angle ofattack 128 with anairflow direction 129 of theairflow 58 through the inlet 60 (FIG. 2 ) of thenacelle 50. Notably, for the embodiment depicted, theairflow direction 129 is substantially parallel to the axial direction A and thelongitudinal axis 12 of theouter nacelle 50 of theturbofan engine 10. For the embodiment depicted, the angle ofattack 128 at the location depicted along thespan 106 of the part spaninlet guide vanes 100 is at least approximately five degrees and up to approximately thirty-five degrees. For example, in certain embodiments, the angle ofattack 128 at the location depicted along thespan 106 of the part spaninlet guide vane 100 may be between about ten degrees and about thirty degrees, such as between about fifteen degrees and about twenty-five degrees. - Additionally, the part span
inlet guide vane 100, at the location depicted along the span 106 (FIG. 2 ) of the part spaninlet guide vane 100 defines alocal swirl angle 130 at the trailingedge 110. The “swirl angle” at the trailingedge 110 of the part spaninlet guide vane 100, as used herein, refers to an angle between theairflow direction 129 of theairflow 58 through the inlet 60 (FIG. 2 ) of the nacelle 50 (FIG. 2 ) and areference line 132 defined by a trailing edge section of thepressure side 120 of the part spaninlet guide vane 100. More specifically, thereference line 132 is defined by the aft twenty percent of thepressure side 120, as measured along thechord line 126. Notably, when the aft twenty percent thepressure side 120 defines a curve, thereference line 132 may be a straight-line average fit of such curve (e.g., using least mean squares). - Further, it will be appreciated, that a
maximum swirl angle 130 refers to thehighest swirl angle 130 along the span 106 (FIG. 2 ) of the part spaninlet guide vane 100. For the embodiment depicted, themaximum swirl angle 130 is defined proximate the radially outer end 102 (FIG. 2 ) of the part span inlet guide vane 100 (e.g., at the outer ten percent of thespan 106 of the part span inlet guide vanes 100), as is represented by the cross-section depicted inFIG. 6 . For the embodiment depicted, themaximum swirl angle 130 of each part spaninlet guide vane 100 at the trailingedge 110 is between approximately five degrees and approximately thirty-five degrees. For example, in certain exemplary embodiments, themaximum swirl angle 130 of each part spaninlet guide vane 100 at the trailingedge 110 may be between twelve degrees and twenty-five degrees. - Moreover, it should be appreciated that for the embodiment of
FIG. 2 , thelocal swirl angle 130 increases from the radially inner end 104 (FIG. 2 ) to the radially outer end 102 (FIG. 2 ) of each part spaninlet guide vane 100. For example, referring now also toFIG. 7 , a cross-sectional view of a part spaninlet guide vane 100 at a location radially inward from the cross-section viewed inFIG. 6 , as indicated by Line 7-7 inFIG. 2 , is provided. As is depicted inFIG. 7 , and as stated above, the part spaninlet guide vane 100 defines thepressure side 120, thesuction side 122, theleading edge 108, the trailingedge 110, thecamber line 124, andchord line 126. Further, the second angle or angle ofattack 128 defined by thechord line 126 and theairflow direction 129 of theairflow 58 through theinlet 60 of thenacelle 50 at the location along thespan 106 depicted inFIG. 7 is less than the angle ofattack 128 at the location along thespan 106 depicted inFIG. 6 (e.g., may be at least about twenty percent less, such as at least about fifty percent less, such as up to about one hundred percent less). Additionally, the part spaninlet guide vane 100 defines alocal swirl angle 130 at the trailingedge 110 at the location along thespan 106 of the part spaninlet guide vane 100 proximate theinner end 104, as depicted inFIG. 7 . As stated above, thelocal swirl angle 130 increases from the radiallyinner end 104 to the radiallyouter end 102 of each part span inlet guide vanes 100. Accordingly, thelocal swirl angle 130 proximate the outer end 102 (seeFIG. 6 ) is greater than thelocal swirl angle 130 proximate the radially inner end 104 (seeFIG. 7 ; e.g., the radially inner ten percent of the span 106). For example, thelocal swirl angle 130 may approach zero degrees (e.g., may be less than about five degrees, such as less than about two degrees) at the radiallyinner end 104. - Notably, including part span
inlet guide vanes 100 of such a configuration may reduce an amount of turbulence at the radially inner end 104 (FIG. 2 ) of each respective part spaninlet guide vane 100. Additionally, such a configuration may provide a desired amount of pre-swirl at the radially outer ends of the plurality of fan blades 40 (FIG. 2 ) of the fan 38 (FIG. 2 ) (where the speed of thefan blades 40 is the greatest) to provide a desired reduction in flow separation and/or shock losses that may otherwise occur due to a relatively high speed of the plurality offan blades 40 at the fan tips during operation of the turbofan engine 10 (FIG. 2 ). - Referring now to
FIG. 8 , an axial view of theinlet 60 to theturbofan engine 10 ofFIGS. 1 and 2 , according to another exemplary embodiment, is provided. Referring toFIG. 8 , in another exemplary embodiment, the plurality of part spaninlet guide vanes 100 are angled at different first angles A1 with respect to the radius NR of theouter nacelle 50. - The
outer nacelle 50 includes atop portion 210, abottom portion 212, afirst side portion 214, and asecond side portion 216. In an exemplary embodiment, a first portion, e.g., thetop portion 210, of the part spaninlet guide vanes 100 are angled at a first angle A1 with respect to the radius NR of theouter nacelle 50. For example, the part spaninlet guide vanes 100 are angled at a first angle A1 of five degrees or seven degrees with respect to the radius NR of theouter nacelle 50 at thetop portion 210. Furthermore, a second portion, e.g., thefirst side portion 214, of the part spaninlet guide vanes 100 are angled at a different angle with respect to the radius NR of theouter nacelle 50. For example, the part spaninlet guide vanes 100 are angled at a first angle A1 of ten degrees or fifteen degrees with respect to the radius NR of theouter nacelle 50 at thefirst side portion 214. In such exemplary embodiments, the plurality of part spaninlet guide vanes 100 are angled at different first angles A1 with respect to the radius NR of theouter nacelle 50 along the circumferential direction C. - In such an exemplary embodiment, circumferential variation in the angle of tilt, e.g., the first angle A1 with respect to the radius NR of the
outer nacelle 50, can address many issues including cross winds, high angle of attack maneuvers such as takeoff, and engine installation on one side of an aircraft or another. For example, cross winds are more likely to affect the 3 or 9 o'clock positions on the engine, e.g., thesecond side portion 216 and thefirst side portion 214, which may make a different angle of tilt more desirable at the 3/9 o'clock positions than at the 6/12 o'clock positions on the engine, e.g., thebottom portion 212 and thetop portion 210. In addition, installation on one side of the aircraft or the other may make cross wind effects more pronounced on the side of the engine that is further from the fuselage. Likewise, a high angle of attack may make differing tilt angles at the 6/12 o'clock positions desirable, e.g., thebottom portion 212 and thetop portion 210. For these reasons, it is contemplated that the first angles A1 with respect to the radius NR of theouter nacelle 50 may vary for a particular application and may be different at thetop portion 210, thebottom portion 212, thefirst side portion 214, and/or thesecond side portion 216 of theouter nacelle 50. - Referring now to
FIG. 9 , an axial view of theinlet 60 to theturbofan engine 10 ofFIGS. 1 and 2 , according to another exemplary embodiment, is provided. Referring toFIG. 9 , in another exemplary embodiment, the plurality of part spaninlet guide vanes 100 are angled at different first angles A1 with respect to the radius NR of theouter nacelle 50. - The
outer nacelle 50 includes atop portion 210, abottom portion 212, afirst side portion 214, and asecond side portion 216. In an exemplary embodiment, the part spaninlet guide vanes 100 are angled at different first angles A1 with respect to the radius NR of theouter nacelle 50 at thetop portion 210 and thebottom portion 212. For example, thetop portion 210 of the part spaninlet guide vanes 100 are angled at a first angle A1 with respect to the radius NR of theouter nacelle 50. For example, the part spaninlet guide vanes 100 are angled at a first angle A1 of five degrees or seven degrees with respect to the radius NR of theouter nacelle 50 at thetop portion 210. Furthermore, thebottom portion 212 of the part spaninlet guide vanes 100 are angled at a different angle with respect to the radius NR of theouter nacelle 50. For example, the part spaninlet guide vanes 100 are angled at a first angle A1 of ten degrees with respect to the radius NR of theouter nacelle 50 at thebottom portion 212. In such exemplary embodiments, the plurality of part spaninlet guide vanes 100 are angled at different first angles A1 with respect to the radius NR of theouter nacelle 50 at thetop portion 210 and thebottom portion 212. - In such an exemplary embodiment, circumferential variation in the angle of tilt, e.g., the first angle A1 with respect to the radius NR of the
outer nacelle 50, can address many issues including cross winds, high angle of attack maneuvers such as takeoff, and engine installation on one side of an aircraft or another. For example, cross winds are more likely to affect the 3 or 9 o'clock positions on the engine, e.g., thesecond side portion 216 and thefirst side portion 214, which may make a different angle of tilt more desirable at the 3/9 o'clock positions than at the 6/12 o'clock positions on the engine, e.g., thebottom portion 212 and thetop portion 210. In addition, installation on one side of the aircraft or the other may make cross wind effects more pronounced on the side of the engine that is further from the fuselage. Likewise, a high angle of attack may make differing tilt angles at the 6/12 o'clock positions desirable, e.g., thebottom portion 212 and thetop portion 210. For these reasons, it is contemplated that the first angles A1 with respect to the radius NR of theouter nacelle 50 may vary for a particular application and may be different at thetop portion 210, thebottom portion 212, thefirst side portion 214, and/or thesecond side portion 216 of theouter nacelle 50. - Referring now to
FIG. 10 , an axial view of theinlet 60 to theturbofan engine 10 ofFIGS. 1 and 2 , according to another exemplary embodiment, is provided. Referring toFIG. 10 , in another exemplary embodiment, one of the part spaninlet guide vanes 100 are angled at different angles with respect to the radius NR of theouter nacelle 50 at different portions of the part spaninlet guide vane 100. - For example, the part span
inlet guide vane 100 is angled at a first angle A1 with respect to the radius NR of the outer nacelle 50 afirst location 250 of the part spaninlet guide vane 100 and the part spaninlet guide vane 100 is angled at a second angle A2 with respect to the radius NR of theouter nacelle 50 at asecond location 252 of the part spaninlet guide vane 100. In such an embodiment, the first angle A1 is different than the second angle A2. For example, in the embodiment depicted inFIG. 10 , the first angle A1 is greater than the second angle A2. - In such an exemplary embodiment, varying the angle of tilt, e.g., the first angle A1 and the second angle A2, from the base of the part span
inlet guide vane 100, e.g., the outer end 102 (FIG. 2 ), to the tip of the part spaninlet guide vane 100, e.g., the inner end 104 (FIG. 2 ), can address varying levels of turbulence due to boundary layer effects, as you get closer to/further from the wall 52 (FIG. 2 ) of theouter nacelle 50. The angle of tilt and sweep may also be used to improve the matching of the swirl that is imparted to the incoming air with the varying linear speed of the fan blade 40 (FIG. 2 ) as you get closer to/further from the axis 12 (FIG. 1 ) of theengine 10. The linear speed of the fan blade 40 (FIG. 2 ) is greatest at the tip, and slowest at the root of the fan blade 40 (FIG. 2 ), even though the rotational velocity is equal. - Referring now to
FIG. 11 , an axial view of theinlet 60 to theturbofan engine 10 ofFIGS. 1 and 2 , according to another exemplary embodiment, is provided. Referring toFIG. 11 , in another exemplary embodiment, one of the part spaninlet guide vanes 100 are angled at different angles with respect to the radius NR of theouter nacelle 50 at different locations of the part spaninlet guide vane 100. - For example, the part span
inlet guide vane 100 is angled at a first angle A1 with respect to the radius NR of the outer nacelle 50 afirst location 250 of the part spaninlet guide vane 100 and the part spaninlet guide vane 100 is angled at a second angle A2 with respect to the radius NR of theouter nacelle 50 at asecond location 252 of the part spaninlet guide vane 100. In such an embodiment, the first angle A1 is different than the second angle A2. For example, in the embodiment depicted inFIG. 11 , the first angle A1 is less than the second angle A2. - In such an exemplary embodiment, varying the angle of tilt, e.g., the first angle A1 and the second angle A2, from the base of the part span
inlet guide vane 100, e.g., the outer end 102 (FIG. 2 ), to the tip of the part spaninlet guide vane 100, e.g., the inner end 104 (FIG. 2 ), can address varying levels of turbulence due to boundary layer effects, as you get closer to/further from the wall 52 (FIG. 2 ) of theouter nacelle 50. The angle of tilt and sweep may also be used to improve the matching of the swirl that is imparted to the incoming air with the varying linear speed of the fan blade 40 (FIG. 2 ) as you get closer to/further from the axis 12 (FIG. 1 ) of theengine 10. The linear speed of the fan blade 40 (FIG. 2 ) is greatest at the tip, and slowest at the root of the fan blade 40 (FIG. 2 ), even though the rotational velocity is equal. - 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, the nacelle defining a radius and a longitudinal axis; and an inlet pre-swirl vane located upstream of the plurality of fan blades and defining a chord, the inlet pre-swirl vane coupled to the nacelle, wherein the inlet pre-swirl vane is angled at a first angle with respect to the radius of the nacelle, and wherein the chord of the inlet pre-swirl vane is angled at a second angle with respect to the longitudinal axis of the nacelle.
- The turbofan engine of any preceding clause, wherein the inlet pre-swirl vane is angled at the first angle with respect to the radius of the nacelle in a clockwise direction from an inlet of the nacelle.
- The turbofan engine of any preceding clause, wherein the inlet pre-swirl vane is angled at the first angle with respect to the radius of the nacelle in a counterclockwise direction from an inlet of the nacelle.
- The turbofan engine of any preceding clause, wherein the first angle is between approximately 2 degrees and approximately 45 degrees.
- The turbofan engine of any preceding clause, wherein the second angle is between approximately 5 degrees and approximately 35 degrees.
- The turbofan engine of any preceding clause, wherein the inlet pre-swirl vane is one of a plurality of part span inlet guide vanes extending from the nacelle upstream of the plurality of fan blades and aft of an inlet of the nacelle.
- The turbofan engine of any preceding clause, wherein each of the plurality of part span inlet guide vanes is angled at the same first angle with respect to the radius of the nacelle.
- The turbofan engine of any preceding clause, wherein a first portion of each of the plurality of part span inlet guide vanes is angled at the first angle with respect to the radius of the nacelle, wherein a second portion of each of the plurality of part span inlet guide vanes is angled at a third angle with respect to the radius of the nacelle, and wherein the first angle is different than the third angle.
- The turbofan engine of any preceding clause, wherein the nacelle includes a top portion, a bottom portion, a first side portion, and a second side portion, wherein a first portion of each of the plurality of part span inlet guide vanes is angled at the first angle with respect to the radius of the nacelle at the top portion, wherein a second portion of each of the plurality of part span inlet guide vanes is angled at a third angle with respect to the radius of the nacelle at the bottom portion, and wherein the first angle is different than the third angle.
- A nacelle assembly for a turbofan engine, the turbofan engine comprising a fan including a plurality of fan blades, the nacelle assembly comprising: a nacelle surrounding and at least partially enclosing the fan, the nacelle defining a radius and a longitudinal axis; and an inlet pre-swirl vane located upstream of the plurality of fan blades and defining a chord, the inlet pre-swirl vane coupled to the nacelle, wherein the inlet pre-swirl vane is angled at a first angle with respect to the radius of the nacelle, and wherein the chord of the inlet pre-swirl vane is angled at a second angle with respect to the longitudinal axis of the nacelle.
- The nacelle assembly of any preceding clause, wherein the inlet pre-swirl vane is angled at the first angle with respect to the radius of the nacelle in a clockwise direction from an inlet of the nacelle.
- The nacelle assembly of any preceding clause, wherein the inlet pre-swirl vane is angled at the first angle with respect to the radius of the nacelle in a counterclockwise direction from an inlet of the nacelle.
- The nacelle assembly of any preceding clause, wherein the first angle is between approximately 2 degrees and approximately 45 degrees.
- The nacelle assembly of any preceding clause, wherein the second angle is between approximately 5 degrees and approximately 35 degrees.
- The nacelle assembly of any preceding clause, wherein the inlet pre-swirl vane is one of a plurality of part span inlet guide vanes extending from the nacelle upstream of the plurality of fan blades and aft of an inlet of the nacelle.
- The nacelle assembly of any preceding clause, wherein each of the plurality of part span inlet guide vanes is angled at the same first angle with respect to the radius of the nacelle.
- The nacelle assembly of any preceding clause, wherein a first portion of each of the plurality of part span inlet guide vanes is angled at the first angle with respect to the radius of the nacelle, wherein a second portion of each of the plurality of part span inlet guide vanes is angled at a third angle with respect to the radius of the nacelle, and wherein the first angle is different than the third angle.
- The nacelle assembly of any preceding clause, wherein the nacelle includes a top portion, a bottom portion, a first side portion, and a second side portion, wherein a first portion of each of the plurality of part span inlet guide vanes is angled at the first angle with respect to the radius of the nacelle at the top portion, wherein a second portion of each of the plurality of part span inlet guide vanes is angled at a third angle with respect to the radius of the nacelle at the bottom portion, and wherein the first angle is different than the third angle.
- A nacelle assembly for a turbofan engine, the turbofan engine comprising a fan including a plurality of fan blades, the nacelle assembly comprising: a nacelle surrounding and at least partially enclosing the fan, the nacelle defining a radius; and an inlet pre-swirl vane located upstream of the plurality of fan blades, the inlet pre-swirl vane coupled to the nacelle, wherein the inlet pre-swirl vane is angled at a first angle with respect to the radius of the nacelle at a first location of the inlet pre-swirl vane, wherein the inlet pre-swirl vane is angled at a second angle with respect to the radius of the nacelle at a second location of the inlet pre-swirl vane, and wherein the first angle is different than the second angle.
- The nacelle assembly of any preceding clause, wherein the nacelle defines a longitudinal axis, wherein the inlet pre-swirl vane defines a chord, and wherein the chord of the inlet pre-swirl vane is angled at a third angle with respect to the longitudinal axis of the nacelle.
- 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 (22)
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US17/676,454 US20230265862A1 (en) | 2022-02-21 | 2022-02-21 | Turbofan engine having angled inlet pre-swirl vanes |
CN202310111481.2A CN116624288A (en) | 2022-02-21 | 2023-02-14 | Turbofan engine with angled inlet pre-swirl vanes |
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US17/676,454 US20230265862A1 (en) | 2022-02-21 | 2022-02-21 | Turbofan engine having angled inlet pre-swirl vanes |
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-
2022
- 2022-02-21 US US17/676,454 patent/US20230265862A1/en active Pending
-
2023
- 2023-02-14 CN CN202310111481.2A patent/CN116624288A/en active Pending
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US6935833B2 (en) * | 2002-02-28 | 2005-08-30 | Mtu Aero Engines Gmbh | Recirculation structure for turbo chargers |
US7575412B2 (en) * | 2002-02-28 | 2009-08-18 | Mtu Aero Engines Gmbh | Anti-stall casing treatment for turbo compressors |
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US20160084265A1 (en) * | 2014-09-23 | 2016-03-24 | Pratt & Whitney Canada Corp. | Gas turbine engine with partial inlet vane |
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