TURBOMACHINE AND METHOD OF ASSEMBLY

Description

CROSS REFERENCE TO RELATED APPLICATIONS

This application is a continuation-in-part of application U.S. Ser. No. 18/654,444, filed May 3, 2024, which is a continuation-in-part of application U.S. Ser. No. 18/511,128, filed Nov. 16, 2023, which is a continuation of application U.S. Ser. No. 18/138,442, filed on Apr. 24, 2023, now U.S. Pat. No. 11,852,161, which is a continuation-in-part of application U.S. Ser. No. 17/986,544, filed on Nov. 14, 2022, now U.S. Patent No. U.S. Pat. No. 11,661,851, the contents of all of which are incorporated herein by reference in their entireties.

FIELD

The present disclosure relates generally to jet engines and, more particularly, to jet engine fans.

BACKGROUND

In one form, a gas turbine engine can include a fan and a core arranged in flow communication with one another. The core generally includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. The fan and the core may be partially surrounded by an outer nacelle. In such approaches, the outer nacelle defines a bypass airflow passage with the core.

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 using one or more fuel nozzles within the combustion section and burned to provide combustion gases. The combustion gases are routed from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the turbine section and is then routed through the exhaust section to atmosphere.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present disclosure, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, with reference to the appended figures, in which:

FIG. 1 is a schematic, cross-sectional view of a gas turbine engine in accordance with exemplary aspects of the present disclosure;

FIG. 2 is a sectional view of a fan blade in accordance with exemplary aspects of the present disclosure;

FIG. 3 shows first example engines arranged on a first plot in accordance with a first performance factor for a fan module according to the present disclosure;

FIG. 4 shows second example engines arranged on a second plot in accordance with a second performance factor for a fan module according to the present disclosure;

FIG. 5 shows third example engines arranged on a third plot in accordance with the first performance factor for a fan module according to the present disclosure;

FIG. 6 shows fourth example engines arranged on a fourth plot in accordance with the second performance factor for a fan module according to the present disclosure;

FIG. 7 is a top planiform view of two adjacent fan blades in accordance with exemplary aspects of the present disclosure;

FIG. 8 illustrates two scaled graphs of airfoil camber and slope thereof versus span for a fan blade illustrated in FIG. 1 and a corresponding preexisting fan blade design; and

FIG. 9 is a side elevational view of a fan blade and a graph of corresponding aerodynamic sweep thereof.

Elements in the figures are illustrated for simplicity and clarity and have not necessarily been drawn to scale. For example, the dimensions and/or relative positioning of some of the elements in the figures may be exaggerated relative to other elements to help to improve understanding of variations of the present disclosure. Also, common but well-understood elements that are useful or necessary in a commercially feasible embodiment are often not depicted in order to facilitate a less obstructed view of these variations of the present disclosure. Certain actions and/or steps may be described or depicted in a particular order of occurrence while those skilled in the art will understand that such specificity with respect to sequence is not actually required.

DETAILED DESCRIPTION

Aspects and advantages of the present disclosure will be set forth in part in the following description or may be learned through practice 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.

As used herein, the terms “first” and “second” 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 “upstream” and “downstream” refer to the relative direction with respect to a flow in a pathway. For example, with respect to a fluid flow, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows.

The terms “radial” or “radially” refer to a direction away from a common center. For example, in the overall context of a turbine engine, radial refers to a direction along a ray extending between a center longitudinal axis of the engine and an outer engine circumference.

The term “composite,” as used herein is, refers to a material that includes non-metallic elements or materials. As used herein, a “composite component” or “composite material” refers to a structure or a component including any suitable composite material. A composite material can be a combination of at least two or more non-metallic elements or materials. Examples of a composite material can be, but not limited to, a polymer matrix composite (PMC), a ceramic matrix composite (CMC), a metal matrix composite (MMC), carbon fibers, a polymeric resin, a thermoplastic resin, bismaleimide (BMI) materials, polyimide materials, an epoxy resin, glass fibers, and silicon matrix materials. Composite components, such as a composite airfoil, can include several layers or plies of composite material. The layers or plies can vary in stiffness, material, and dimension to achieve the desired composite component or composite portion of a component having a predetermined weight, size, stiffness, and strength. One or more layers of adhesive can be used in forming or coupling composite components. Adhesives can include resin and phenolics, wherein the adhesive can require curing at elevated temperatures or other hardening techniques.

As used herein, “PMC” refers to a class of materials. By way of example, the PMC material is defined in part by a prepreg, which is a reinforcement material pre-impregnated with a polymer matrix material, such as thermoplastic resin. Non-limiting examples of processes for producing thermoplastic prepregs include hot melt pre-pregging in which the fiber reinforcement material is drawn through a molten bath of resin and powder pre-pregging in which a resin is deposited onto the fiber reinforcement material, by way of non-limiting example electrostatically, and then adhered to the fiber, by way of non-limiting example, in an oven or with the assistance of heated rollers. The prepregs can be in the form of unidirectional tapes or woven fabrics, which are then stacked on top of one another to create the number of stacked plies desired for the part. Multiple layers of prepreg may be stacked to the proper thickness and orientation for the composite component and then the resin is cured and solidified to render a fiber reinforced composite part. Resins for matrix materials of PMCs can be generally classified as thermosets or thermoplastics. Thermoplastic resins are generally categorized as polymers that can be repeatedly softened and flowed when heated and hardened when sufficiently cooled due to physical rather than chemical changes. Notable example classes of thermoplastic resins include nylons, thermoplastic polyesters, polyaryletherketones, and polycarbonate resins. Specific example of high performance thermoplastic resins that have been contemplated for use in aerospace applications include, polyetheretherketone (PEEK), polyetherketoneketone (PEKK), polyetherimide (PEI), polyaryletherketone (PAEK), and polyphenylene sulfide (PPS). In contrast, once fully cured into a hard rigid solid, thermoset resins do not undergo significant softening when heated, but instead thermally decompose when sufficiently heated. Notable examples of thermoset resins include epoxy, bismaleimide (BMI), and polyimide resins.

Instead of using a prepreg, in another non-limiting example, with the use of thermoplastic polymers, it is possible to utilize a woven fabric. Woven fabric can include, but is not limited to, dry carbon fibers woven together with thermoplastic polymer fibers or filaments. Non-prepreg braided architectures can be made in a similar fashion. With this approach, it is possible to tailor the fiber volume of the part by dictating the relative concentrations of the thermoplastic fibers and reinforcement fibers that have been woven or braided together. Additionally, different types of reinforcement fibers can be braided or woven together in various concentrations to tailor the properties of the part. For example, glass fibers, carbon fibers, and thermoplastic fibers could all be woven together in various concentrations to tailor the properties of the part. The carbon fibers provide the strength of the system, the glass fibers can be incorporated to enhance the impact properties, which is a design characteristic for parts located near the inlet of the engine, and the thermoplastic fibers provide the binding for the reinforcement fibers.

In yet another non-limiting example, resin transfer molding (RTM) can be used to form at least a portion of a composite component. Generally, RTM includes the application of dry fibers or matrix material to a mold or cavity. The dry fibers or matrix material can include prepreg, braided material, woven material, or any combination thereof.

Resin can be pumped into or otherwise provided to the mold or cavity to impregnate the dry fibers or matrix material. The combination of the impregnated fibers or matrix material and the resin are then cured and removed from the mold. When removed from the mold, the composite component can require post-curing processing.

It is contemplated that RTM can be a vacuum assisted process. That is, the air from the cavity or mold can be removed and replaced by the resin prior to heating or curing. It is further contemplated that the placement of the dry fibers or matrix material can be manual or automated.

The dry fibers or matrix material can be contoured to shape the composite component or direct the resin. Optionally, additional layers or reinforcing layers of material differing from the dry fiber or matrix material can also be included or added prior to heating or curing.

The singular forms “a”, “an”, and “the” include plural references unless the context clearly dictates otherwise.

The inventors have sought to maximize efficiency of turbine engines during in-flight propulsion of an aircraft, and correspondingly reduce fuel consumption. In particular, the inventors were focused on how the fan of a ducted turbine engine can be improved. The inventors, in consideration of several different engine architectures proposed, considered how the fan module would need to change to achieve mission requirements, and/or how the fan module could improve upon an existing engine efficiency and/or fuel consumption. The inventors looked at several engine architectures, then determined how the number of fan blades utilized with a fan, average chord width of the fan blades, diameter of the fan blade, fan pressure ratio, fan tip speed, and hub-to-tip ratio affect engine efficiency and/or fuel consumption.

The inventors found that in some engines, an excess number of fan blades may add unnecessary cost to the engine design without appreciable benefit, and may also add unnecessary weight to an aircraft, thereby reducing overall fuel efficiency (e.g., due to increased fuel burn). A reduction in fan blade quantity, however, was found to potentially lead to a reduction in total fan blade area desired for efficient propulsion, fan aeromechanical stability and operability, etc. The inventors considered increasing the width or fan chord of the fan blades to achieve a desired fan blade area with a lower fan blade count. Such considerations were found to be of particular interest when the engine had a higher bypass ratio (i.e., lower fan pressure ratio), and when the engine had a lower blade tip speed.

The determination of the fan blade count and average fan chord for achieving a desired efficiency often required a time consuming, iterative process. As explained in greater detail below, after evaluation of numerous turbine engine architectures having different fan blade counts and average fan chords, it was found, unexpectedly, that there exist certain relationships between a fan or fan blade diameter, a fan pressure ratio, and a redline corrected fan tip Mach number of the turbine engine that identify an average fan chord needed to produce improved results in terms of engine efficiency. It was further found, unexpectedly, that there exist certain relationships between turbine engine parameters including a hub-to-tip ratio of the fan, a fan pressure ratio, and a redline corrected fan tip Mach number of the turbine engine that identify a fan blade count needed to produce improved results in terms of engine efficiency.

Various aspects of the present disclosure describe aspects of an aircraft turbine engine characterized in part by an increased average fan chord width and a reduced blade count, which are believed to result in an improved engine aerodynamic efficiency and/or improved fuel efficiency. According to the disclosure, a turbomachine for powering an aircraft in flight comprises an annular casing and a fan disposed inside the annular casing and mounted for rotation about an axial centerline, the fan including fan blades that extend radially outwardly toward the annular casing.

Reference will now be made in detail to embodiments of the disclosure, one or more examples of which are illustrated in the accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the disclosure and, together with the description, serve to explain the principles of the disclosure. 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.

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 turbomachine, more specifically 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 a high-bypass turbofan jet engine 10, referred to herein as “turbofan jet engine 10.” As shown in FIG. 1, the turbofan jet engine 10 defines an axial direction A (extending parallel to a longitudinal centerline 12 provided for reference) and a radial direction R. In general, the turbofan jet engine 10 includes a fan section 14 and a core turbine engine 16 disposed downstream from the fan section 14.

The exemplary core turbine engine 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 booster or low pressure (LP) compressor 22 and a high pressure (HP) compressor 24; a combustion section 26; a turbine section including a high pressure (HP) turbine 28 and a low pressure (LP) turbine 30; and a jet exhaust nozzle section 32. A high pressure (HP) shaft or spool 34 drivingly connects the high pressure turbine 28 to the high pressure compressor 24. A low pressure (LP) shaft or spool 36 drivingly connects the low pressure turbine 30 to the low pressure compressor 22.

Fan blades 40 extend outwardly from disk 42 generally along the radial direction R. For the embodiment depicted, the fan section 14 includes a variable pitch fan 38 having a plurality of fan blades 40 coupled to a disk 42 in a spaced apart manner. One or more of the fan blades 40 may 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 vary the pitch of the fan blades 40, typically collectively in unison. In some approaches, the fan is a fixed pitch fan and actuation member 44 is not present. The fan blades 40, disk 42, and actuation member 44 may be together rotatable about the longitudinal centerline 12 by low pressure spool 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 low pressure spool 36 to a more efficient rotational fan speed. In some approaches, the low pressure spool 36 may directly drive the fan without power gear box 46.

The power gear box 46 can include a plurality of gears, including an input and an output, and may also include one or more intermediate gears disposed between and/or interconnecting the input and the output. The input can comprise a first rotational speed and the output can have a second rotational speed. In some examples, a gear ratio of the first rotational speed to the second rotational speed is equal to or greater than 3.2 and equal to or less than 5.0. The power gear box 46 can comprise various types and/or configurations. In some examples, the power gear box 46 is a single-stage gear box. In other examples, the power gear box 46 is a multi-stage gear box. In some examples, the power gear box 46 is an epicyclic gearbox. In some examples, the power gear box 46 is a non-epicyclic gear box (e.g., a compound gearbox). More particularly, in some instances, the power gear box 46 is an epicyclic gear box configured in a star gear configuration. Star gear configurations comprise a sun gear, a plurality of star gears (which can also be referred to as “planet gears”), and a ring gear. The sun gear is the input and is coupled to the power turbine (e.g., the low-pressure turbine) such that the sun gear and the power turbine rotate at the same rotational speed. The star gears are disposed between and interconnect the sun gear and the ring gear. The star gears are rotatably coupled to a fixed carrier. As such, the star gears can rotate about their respective axes but cannot collectively orbit relative to the sun gear or the ring gear. As another example, the power gear box 46 is an epicyclic gear box configured in a planet gear configuration. Planet gear configurations comprise a sun gear, a plurality of planet gears, and a ring gear. The sun gear is the input and is coupled to the power turbine. The planet gears are disposed between and interconnect the sun gear and the ring gear. The planet gears are rotatably coupled to a rotatable carrier. As such, the planet gears can rotate about their respective axes and also collectively rotate together with the carrier relative to the sun gear and the ring gear. The carrier is the output and is coupled to the fan assembly. The ring gear is fixed from rotation.

Referring still to the exemplary embodiment of FIG. 1, the disk 42 is covered by rotatable front hub 48 aerodynamically contoured to promote an airflow through the plurality of fan blades 40. Additionally, the exemplary fan section 14 includes an annular fan casing or outer nacelle 50 that circumferentially surrounds the variable pitch fan 38 and/or at least a portion of the core turbine engine 16. It should be appreciated that the outer nacelle 50 may be configured to be supported relative to the core turbine engine 16 by a plurality of circumferentially spaced outlet guide vanes 52. Moreover, a downstream section 54 of the outer nacelle 50 may extend over an outer portion of the core turbine engine 16 so as to define a bypass airflow passage 56 therebetween.

During operation of the turbofan jet engine 10, a volume of air 58 enters the turbofan jet engine 10 through an associated inlet 60 of the outer nacelle 50 and/or fan section 14. As the volume of air 58 passes across the fan blades 40, a first portion 62 of the air 58 as indicated by arrow 62 is directed or routed into the bypass airflow passage 56 and a second portion 64 of the air 58 as indicated by arrow 64 is directed or routed into the low pressure compressor 22. The ratio between the first portion 62 of air 58 and the second portion 64 of air 58 is commonly known as a bypass ratio. The pressure of the second portion 64 of air 58 is then increased as it is routed through the high pressure (HP) compressor 24 and into the combustion section 26, where it is mixed with fuel and burned to provide combustion gases 66. Subsequently, the combustion gases 66 are routed through the high pressure turbine 28 and the low pressure turbine 30, where a portion of thermal and/or kinetic energy from the combustion gases 66 is extracted.

The combustion gases 66 are then routed through the jet exhaust nozzle section 32 of the core turbine engine 16 to provide propulsive thrust. Simultaneously, the pressure of the first portion 62 of air 58 is substantially increased as the first portion 62 of air 58 is routed through the bypass airflow passage 56 before it is exhausted from a fan nozzle exhaust section 76 of the turbofan jet engine 10, also providing propulsive thrust.

It should be appreciated, however, that the turbofan jet engine 10 depicted in FIG. 1 is by way of example only, and that in other exemplary embodiments, aspects of the present disclosure may additionally, or alternatively, be applied to any other suitable gas turbine engine. For example, in other exemplary embodiments, the turbofan jet engine 10 may instead be any other suitable aeronautical gas turbine engine, such as a turbojet engine, turboshaft engine, turboprop engine, etc. Additionally, in still other exemplary embodiments, the turbofan jet engine 10 may include or be operably connected to any other suitable accessory systems. Additionally, or alternatively, the exemplary turbofan jet engine 10 may not include or be operably connected to one or more of the accessory systems discussed above.

The fan blades 40 of the jet engine 10 may be made from a PMC material with metal leading edges to protect the airfoil from foreign objects, such as bird strikes. A polymer matrix composite (PMC) material for the airfoil can be more durable and/or exhibit improved performance when the airfoil is subjected to flutter effects during operation. In some embodiments, engines with fewer fan blades (e.g., less than 25 fan blades) and wider chords (c), such as engines having a blade count (BC) from 14 and 18, or 16 to 20 fan blades and ratios of chord to diameter (c/D) of greater than 0.17, or greater than 0.19, and less than 0.3 (e.g., less than 0.21) have the fan blade airfoil made from a PMC material with metal leading edge.

FIG. 2 is a sectional view of a fan blade 40 viewed radially (e.g., towards the rotation axis). A first axis 100 is parallel to the axial direction A of FIG. 1, and a second axis 102 is parallel to the circumferential direction θ.

Fan blade 40 includes a low-pressure surface 110 and an opposite high-pressure surface 112 that each extend between a proximal end 40a and a distal end 40b of the fan blade 40 (shown in FIG. 1). Fan blade 40 further includes a leading edge 114 and a trailing edge 116.

The low-pressure surface 110, high-pressure surface 112, leading edge 114, and trailing edge 116 form a profile 118 of the fan blade 40. The profile 118 defines a mean camber 120 that extends from the leading edge 114 to the trailing edge 116 and that is equidistant from the low-pressure surface 110 and the high-pressure surface 112.

The profile 118 further defines a local chord 122 (relative to a specific cross section through the blade) that represents a straight-line distance from the leading edge 114 to the trailing edge 116.

In some approaches, a fan blade 40 may have a profile 118 that varies along a radial height of the fan blade 40 between the proximal end 40a and the distal end 40b. For example, in some fan blade designs, a distance between the leading edge 114 and the trailing edge 116 may be greater at the proximal end 40a of the fan blade 40 than at the distal end 40b. As such, the length of the chord line 122 may vary along the radial height of the fan blade 40. In this way, an average chord line length may be derived for the fan blade that accounts for the variation in chord line lengths 122 along the radial height of the fan blade 40.

As mentioned earlier, the inventors have discovered relationships between timescales that include a fan pressure ratio, fan diameter, and corrected fan tip Mach number during the course of improving upon the fan module portion of various engine architectures. More particularly, and as discussed in greater detail below, the inventors have discovered relationships between ratio of axial flow timescales to rotation timescales, and suitable parameters for implementing those relationships with an engine.

The aircraft turbine engine architectures developed by the inventors include as major components a fan module and an engine core. The core includes one or more compressor stages and turbine stages. Compressor stages typically include high pressure and low pressure compressor stages, and turbines similarly include high and low pressure stages. The fan module that provides for an improved efficiency is not independent of these other parts of the engine, because there is always a trade benefit when one part is improved or modified. Improved efficiency brought by the fan can be in terms of a reduction in weight, lower installed drag, load balancing or management (dynamic or static loading), aerodynamic efficiency through the fan duct/interaction of fan to output guide vanes, and other factors. In an effort to improve upon what the fan can deliver (positive benefit of fan design) there often times need to be sacrifices in other parts of the engine (negative benefit of fan design). Or the benefits of a new fan design when viewed independent of a particular core design or airframe requirement, often times requires revision or is unrealistic given the impact that such a fan design will have on other parts of the engine, e.g., compressor operating margin, balance of a fan and output guide vanes (OGV) along with a power gearbox, and location of a low pressure compressor (packaging impacts). The teachings described herein are also applicable to other engine architectures such as electrically-driven fans (which may or may not include a turbine) and hybrid electrically-driven fans (e.g., distributed electric propulsion systems in which a gas turbine drives multiple fans).

The inventors, proceeding in the manner of designing improved fan modules, accounting for the trade-offs between fan module improvements and other potentially negative or limitations on fan module design, unexpectedly found certain relationships that define an improved fan design, now described in detail.

In one aspect, the inventors have discovered a relationship between an average fan chord “c”, a fan diameter “D” (e.g., a tip-to-tip dimension of the fan), a fan pressure ratio “FPR” of a fan, and a corrected redline fan tip Mach number

“ M tip , c ( RL ) ”

according to the below relationship, referred to herein as the First Performance Factor (“FPF”) for a fan module:

FPF = [ c 0.15 · D ] / [ [ FPR - 1 0.4 ] / M tip , c ( RL ) ] - 1 .23 ( 1 ) m 1 · [ M tip , c ( RL ) - 1 . 1 ] + 6 > FPF > m 1 · [ M tip , c ( RL ) - 1 . 1 ] + Δ ⁢ y 1 ( 2 )

The ratio of average fan chord “c” to fan diameter “D” is a nondimensionalized chord width ratio greater than 0.1 (e.g., greater than 0.15, greater than 0.17, or greater than 0.19), and less than 0.3 (e.g., less than 0.21).

As used herein, the “fan pressure ratio” (FPR) refers to a ratio of a stagnation pressure immediately downstream of the plurality of outlet guide vanes 52 during operation of the fan 38 to a stagnation pressure immediately upstream of the plurality of fan blades 40 during the operation of the fan 38. The “√{square root over (FPR−1)}” portion of the average fan chord relationship may be utilized as a surrogate for referencing a proportionality to the increase in axial flow velocity through the fan. The fan pressure ratio is greater than 1.2 (e.g., greater than 1.3), and less than 1.5 (e.g., less than 1.45, less than 1.42, or less than 1.4).

As used herein,

“ M tip , c ( RL ) ”

is a corrected fan tip Mach number at redline (e.g., maximum permissible rotational speed of the fan at a redline shaft speed, which is either directly coupled to the fan or through a reduction gearbox). “Fan tip speed” refers to a linear speed of an outer tip of a fan blade 40 during operation of the fan 38. “Corrected fan tip speed” (referred to as “U_tip,c”) may be provided, for example, as ft/sec divided by an industry standard temperature correction. In an example approach, U_tip,c may be less than 1,500 ft/sec (e.g., less than 1,250 ft/sec or less than 1,100 ft/sec), and greater than 500 ft/sec. “Corrected fan tip Mach number” refers to a nondimensionalized value obtained by dividing U_tip,c by the generally accepted speed of sound at standard day sea level atmospheric conditions (i.e., 1,116.45 ft/sec). As such,

M tip , c ( RL )

may be less than 1.34 (e.g., less than 1.12 or less than 0.99), and greater than 0.45.

FPF, as defined in (1), may be thought of as representing a ratio of speeds. When considered with the normalized chord width “c/D,” FPF may be thought of as a correlation of timescales of the blade rotation with the time taken for a flow particle to traverse a fan average chord length when the engine is operating at static conditions.

Referring to the inequality defined in (2) and to the plot of FIG. 3, example engine embodiments are shown having unique FPF values and corresponding redline corrected fan tip Mach number

( M tip , c ( RL ) ) .

FPF increases in value along the Y-axis, while the X-axis represents left-to-right increasing redline corrected fan tip Mach number

( M tip , c ( RL ) ) .

FIG. 3 also shows a first line 200 and a second line 202 that is offset from the first line 200 along the Y-axis. The first and second lines 200, 202 are defined by the

“ m 1 · [ M tip , c ( RL ) - 1.1 ] + Δ ⁢ y 1 ”

portion of inequality (2). As used herein, “m₁” refers to a slope of a line 200, 202, “1.1” refers to a reference corrected redline tip Mach number at which Y-intercept is defined in the FPF, and Δy₁ refers an offset from the Y-intercept along the Y-axis.

As shown in FIG. 3, the first and second lines 200, 202 are piecewise linear dividing curves; i.e., the first and second lines 200, 202 have different slopes “m₁” depending on the

M tip , c ( RL )

along the X-axis. More particularly, when the value of

M tip , c ( RL )

is equal to or greater than 1.1, the first and second lines 200, 202 have slopes “m₁” equal to 0.87. When the value of

M tip , c ( RL )

is less than 1.1, the first and second lines 200, 202 have slopes “m₁” equal to 3.34. While depicted as piecewise linear dividing curves, the low-speed scaling is actually nonlinear and there are advantages to lower c/D designs toward the lower portion of the plot of FIG. 3 associated with lower FPR and lower

M tip , c ( RL ) .

As discussed, Δy₁ refers an offset from the Y-intercept along the Y-axis. The Δy₁ value can be 0.0125, 0.04, 0.07, 0.1, or 0.2, or can vary between 0 and 6, 0 and 0.0125, 0.0125 and 0.04, 0.04 and 0.07, 0.07 and 0.1, 0.1 and 0.2, or a value greater than 0.2 and less than 6.

FIG. 3 shows eight example engine embodiments, of which engines 210, 212, 214, 216 may be referred to as low speed engine designs (as indicated by subplot area 218), and engines 220, 222, 224, 226 may be referred to as high speed engine designs (as indicated by subplot area 228). Each of the engines 210, 212, 214, 216, 220, 222, 224, 226 have a gear ratio in a range equal to or greater than 3.2 and less than or equal to 5.0.

As represented by the FPF, which indicates a particular fan chord relationship, the inventors discovered a limited or narrowed selection of average fan chords that uniquely take into consideration other factors associated with the fan and engine type. For instance, the inventors determined that an engine having an FPF value for a given

M tip , c ( RL )

value above line 200 (within plot area 240) may allow for relatively wider chord widths as compared to engines having an FPF value for a given

M tip , c ( RL )

value below line 200 (within plot area 242). In this way, engines 214, 216, 224, and 226 may provide advantages over engines 210, 212, 220, and 222, such as a reduced fan blade count (discussed in greater detail below), increased aeromechanical stability and reduced fan lift coefficient CL during takeoff of the aircraft. In some instances, such advantages may become more pronounced as FPF increases and

M tip , c ( RL )

value decreases (for next generation ultra-high bypass ratio engines for instance). For example, the improvement in engine performance based on the redline tip Mach number may have FPF values greater than

m 1 · [ M tip , c ( RL ) - 1.1 ] + 0 . 0 ⁢ 125 , greater ⁢ than ⁢ m 1 · [ M tip , c ( RL ) - 1.1 ] + 0.04 , 
 greater ⁢ than ⁢ m 1 · [ M tip , c ( RL ) - 1.1 ] + 0.07 , greater ⁢ than ⁢ m 1 · [ M tip , c ( RL ) - 1.1 ] + 0 .1 ,

or greater than

m 1 · [ M tip , c ( RL ) - 1.1 ] + 0 . 2

(these other examples are schematically represented by the phantom line 202).

In another aspect, the inventors have discovered a relationship between a fan blade count “BC”, a hub-to-tip ratio of a fan “HTR”, a fan pressure ratio “FPR” of a fan, and a corrected redline fan tip Mach number

“ M tip , c ( RL ) ”

according to the below relationship, referred to herein as the Second Performance Factor (“SPF”) for a fan module:

SPF = π 4 ⁢ ( 1 - HTR 2 ) / ( BC 20 ) / ( FPR - 1 4. ) / M tip , c ( RL ) - 0.97 ( 3 ) m 2 · [ M tip , c ( RL ) - 1.1 ] + 1.5 > SPF > m 2 · [ M tip , c ( RL ) - 1.1 ] + Δ ⁢ y 2 ( 4 )

Regarding the hub-to-tip ratio “HTR,” a fan blade defines a hub radius (R_hub), which is the radius of the leading edge at the hub relative to a centerline of the fan, and a tip radius (R_tip), which is the radius of the leading edge at a tip of the fan blade relative to the centerline of the fan. HTR is the ratio of the hub radius to the tip radius (R_hub/R_tip). The ratio is greater than 0.1 and less than 0.5 (e.g., less than 0.275, less than 0.25, or less than 0.225).

Blade count “BC” corresponds to the number of fan blades circumferentially arranged about the fan hub. The blade count is between 10 fan blades and 40 fan blades. In certain example approaches, the blade count is less than or equal to 18 fan blades (e.g., 16 or fewer fan blades).

“FPR” and

“ M tip , c ( RL ) ”

refer to a tan pressure ratio and a redline corrected fan tip Mach number, respectively, as discussed with respect to the average fan chord relationship above. In this way, the values of one or more of the FPR and

“ M tip , c ( RL ) ”

may be the same as those discussed with respect to the average fan chord relationship.

Referring to the inequality defined in (4) and to the plot of FIG. 4, example engine embodiments are shown having unique SPF values and corresponding redline corrected fan tip Mach number

( M tip , c ( RL ) ) .

SPF increases in value along the Y-axis, while the X-axis represents left-to-right increasing redline corrected fan tip Mach number

( M tip , c ( RL ) ) .

FIG. 4 also shows a first line 300 and a second line 302 that is offset from the first line 300 along the Y-axis. The first and second lines 300, 302 are defined by the

“ m 2 · [ M tip , c ( RL ) - 1 . 1 ] + Δ ⁢ y 2 ”

portion of inequality (4).

As shown in FIG. 4, the first and second lines 300, 302 are piecewise linear dividing curves; i.e., the first and second lines 300, 302 have different slopes “m₂” depending on the

M tip , c ( RL )

along the X-axis. More particularly, when the value of

M tip , c ( RL )

is equal to or greater than 1.1, the first and second lines 300, 302 have slopes “m₂” equal to 0.41. When the value of

M tip , c ( RL )

is less than 1.1, the first and second lines 300, 302 have slopes “m₂” equal to 0.55.

As used herein, “m₂” refers to a slope of a line 300, 302, which as shown, is equal to 1. “1.1” refers to a reference corrected redline tip Mach number at which the Y-intercept is defined, and Δy₂ refers an offset from the Y-intercept along the Y-axis. The Δy₂ value can be 0.0075, 0.01, 0.02, 0.024, 0.037, 0.04, or 0.06, or can vary between 0 and 1.5, 0 and 0.0075, 0.0075 and 0.01, 0.01 and 0.2, 0.2 and 0.024, 0.024 and 0.037, 0.037 and 0.04, 0.04 and 0.6, or a value greater than 0.6 and less than 1.5.

FIG. 4 shows eight example engine embodiments, of which engines 310, 312, 314, 316 may be referred to as low speed engine designs (as indicated by subplot area 318), and engines 320, 322, 324, 326 may be referred to as high speed engine designs (as indicated by subplot area 328). Each of the engines 310, 312, 314, 316, 320, 322, 324, 326 have a gear ratio a range equal to or greater than 3.2 and less than or equal to 5.0.

As represented by the SPF, which indicates a particular fan blade count relationship, the inventors discovered a limited or narrowed selection of fan blade count that uniquely take into consideration other factors associated with the fan and engine type. For instance, the inventors determined that an engine having an SPF value for a given

M tip , c ( RL )

value above line 300 (within plot area 340) may allow for reduced fan blade counts as compared to engines having an SPF value for a given

M tip , c ( RL )

value below line 300 (within plot area 342). In this way, engines 314, 316, 324, and 326 may provide advantages over engines 310, 312, 320, and 322, such as a reduced cost and weight. In some instances, such advantages may become more pronounced as the SPF value increases and the

M tip , c ( RL )

value decreases (tor next generation ultra-high bypass ratio engines for instance). For example, the improvement in engine performance based on the redline tip Mach number may have SPF values greater than

m 2 · [ M tip , c ( RL ) - 1.1 ] + 0 . 0 ⁢ 0 ⁢ 75 , greater ⁢ than ⁢ m 2 · [ M tip , c ( RL ) - 1.1 ] + 0.01 , greater ⁢ than ⁢ m 2 · [ M tip , c ( RL ) - 1.1 ] + 0.02 , greater ⁢ than ⁢ m 2 · [ M tip , c ( RL ) - 1.1 ] + 0.024 , greater ⁢ than ⁢ m 2 · [ M tip , c ( RL ) - 1.1 ] + 0.037 , greater ⁢ than ⁢ m 2 · [ M tip , c ( RL ) - 1.1 ] + 0.04 ,

or greater than

m 2 · [ M tip , c ( RL ) - 1 . 1 ] + 0 . 0 ⁢ 6

(these other examples are schematically represented by the phantom line 302).

FIG. 5 shows additional example engine embodiments 402, 404, 406, 408 having First Performance Factor (FPF) values, as similarly described herein. Line 400 is a piecewise linear dividing curve having different slopes “m₁” depending on the

M tip , c ( RL )

along the X-axis. More particularly, when the value of

M tip , c ( RL )

is equal to or greater than 1.1, line 400 has a slope “m₁” equal to 9.43. When the value of

M tip , c ( RL )

is less than 1.1, line 400 has a slope “m” equal to 27.02.

Line 420 corresponds to line 200 of FIG. 3 and is similarly a piecewise linear dividing curve having different slopes “m₂” depending on the

M tip , c ( RL )

along the X-axis. As with FIG. 3, when the value of

M tip , c ( RL )

is equal to or greater than 1.1, the line 420 has a slope “m₂” equal to 0.87. When the value of

M tip , c ( RL )

is less than 1.1, line 420 has a slope “m₂” equal to 3.34.

In this approach, the First Performance Factor (FPF) is as provided:

FPF = [ c 0.15 · D ] ⁢ / [ [ FPR - 1 0.4 ] / M tip , c ( RL ) ] - 1.23 ( 5 ) m 1 · [ M tip , c ( RL ) - 1.1 ] + 9.14 > FPF > m 2 · [ M tip , c ( RL ) - 1.1 ] ( 6 )

The First Performance Factor (FPF) values may be in the range, for example, equal to or greater than −0.8 and equal to or less than 8.4, equal to or greater than 0 and equal to or less than 6, or equal to or greater than 1 and equal to or less than 2.

M tip , c ( RL )

values may be within a range equal to or greater than 0.8 and equal to or less than 1.5, or equal to or greater than 0.9 and equal to or less than 1.4. FPR values may be within a range equal to or greater than 1.2 and equal to or less than 1.6, equal to or greater than 1.3 and equal to or less than 1.5, or equal to or greater than 1.35 and equal to or less than 1.45.

FIG. 6 shows additional example engine embodiments 452, 454, 456, 458 having Second Performance Factor (SPF) values, as similarly described herein. Line 450 is a linear curve having slope “m₃” of 3.17. Line 470 corresponds to line 300 of FIG. 4 and is similarly a piecewise linear dividing curve having different slopes “m₄” depending on the

M tip , c ( RL )

along the X-axis. More particularly, when the value of

M tip , c ( RL )

is equal to or greater than 1.1, line 470 has a slope “m₄” equal to 0.41. When the value of

M tip , c ( RL )

is less than 1.1, line 420 has a slope “m₄” equal to 0.55.

In this approach, the Second Performance Factor (SPF) is as provided:

SPF = π 4 ⁢ ( 1 - HTR 2 ) / ( BC 20 ) / ( FPR - 1 0.4 ) / M tip , c ( RL ) - 0.97 ( 7 ) m 3 · [ M tip , c ( RL ) - 1.1 ] + 2.52 > SPF > m 4 · [ M tip , c ( RL ) - 1.1 ] ( 8 )

The Second Performance Factor (SPF) values may be in the range, for example, equal to or greater than 0.087 and equal to or less than 2.4, or equal to or greater than 1 and equal to or less than 2.

M tip , c ( RL )

values may be within a range equal to or greater than 0.8 and equal to or less than 1.5, or equal to or greater than 0.9 and equal to or less than 1.4. HTR values may be within a range equal to or greater than 0.2 and equal to or less than 0.4, or equal to or greater than 0.25 and equal to or less than 0.35.

The inventors have discovered a relationship between the First Performance Factor and Second Performance Factor described herein. More particularly, the inventors have discovered that engine designs achieving each of the First Performance Factor and Second Performance Factor, while maintaining relatively-constant solidity, provide improved engine performance. Blade solidity is defined as the ratio of the blade chord length to the distance of space between the blades.

Example engine parameters and corresponding First Performance Factors and Second Performance Factors are presented in Table 1 below.

In this way, using fan parameters such as fan pressure ratios, corrected fan tip Mach number, fan diameters, and hub-to-tip ratios, the inventors discovered approaches that utilize the above-described average fan chord relationship to obtain an average chord width, and the above-described fan blade count relationship to obtain a fan blade count. These obtained constraints guide one to select fan chord width, blade count, or both suited for the particularized engine architectures and mission requirements, informed by engine-unique environments and trade-offs in design (as discussed above), which are believed to result in an improved engine.

In another aspect, the FPF and SPF may also be useful as a design tool for down-selecting, or providing a guideline for reducing the number of candidate designs for fan blade counts and average fan chords from which to design a fan module for a particular architecture. In this way, an engine architecture is improved overall by knowing, early in the design process, what constraints or limitations would be imposed by a fan module given the mission objectives.

In another aspect method of assembly is provided. The method includes mounting a fan inside an annular casing for rotation about an axial centerline. The fan including fan blades that extend radially outwardly toward the annular casing. The fan further includes an average fan chord width according to the First Performance Factor (“FPF”) and/or a quantity of fan blades according to the Second Performance Factor (“SPF”) discussed above.

As disclosed herein, fan parameters such as fan pressure ratios, corrected fan tip speeds, fan diameters, and hub-to-tip ratios may be used to select a fan chord width, a blade count, or both to provide a turbine engine having improved engine aerodynamic efficiency and/or improved fuel efficiency. During the course of designing a more efficient turbine engine, it was found that certain fan sweep configurations improve both the flow capacity and part-speed stability of engine performance, thereby improving engine aerodynamic efficiency and/or improved fuel efficiency.

Regarding flow capacity, the inventors have found that a swept fan blade as disclosed herein can be more aerodynamically efficient at high rotational speeds by reducing or delaying the onset of shock waves on the blade surfaces. As the fan operates at transonic speeds, a swept blade according to the disclosure herein allows the relative velocity of the airflow over the blade to remain subsonic for a longer portion of the blade, thus reducing drag and improving the flow capacity.

The inventors have further found that a swept blade described below can also reduce tip losses by effectively reducing the blade's aspect ratio (span to chord length) at the tip. This can lead to a more uniform flow distribution across the blade span and reduce tip vortex strength, which in turn can improve the overall flow capacity of the fan.

Still further, sweep can influence the distribution of aerodynamic loading along the blade span. With the sweep angle provided herein, the inventors have found improved control over loading to achieve a more desirable pressure distribution that maximizes the mass flow rate through the fan for a given fan pressure ratio.

In many instances, turbofans are designed for an operating range from takeoff to cruise to landing, with cruise operation being predominant and for which maximum efficiency and operability is desired. However, part-speed performance must also be considered in good turbofan design and accommodated by the higher camber introduced near the blade tips for the low solidity turbofan design. As discussed, the inventors have further found that aerodynamic sweep as provided herein provides for improved part-speed stability of engine performance.

More particularly, in regard to stall margin, at lower speeds (part-speed), fan blades are more susceptible to stall due to the lower momentum of the incoming air and unfavorable angle of attack on the blade. A sweep as described herein can delay stall by influencing the flow separation characteristics, thus enhancing the operating stability of the fan at various speeds.

Blade sweep can also affect the structural dynamics of the fan blades. A sweep according to the disclosure herein reduces resonance frequencies that occur during part-speed operation, thereby improving the stability and reducing the likelihood of blade vibrations or flutter, which can lead to structural failure.

The fan blade with the inventor's combination of aerodynamic sweep in its leading edge and locally increased camber in the outer panel is effective for increasing hub supercharging of the fan air while maintaining aeromechanical stability. The increased camber of the airfoil outer panel is further effective for delaying flutter and tailors the efficiency characteristics of the fan for improvement particularly at part speed corresponding with aircraft cruise operation.

Furthermore, the inventors have found that the reduction in fan blade number, while maintaining a similar or larger chord to diameter C/D ratio at the airfoil tips, has significant advantages, including an increase in efficiency while maintaining adequate stability and stall margin, as well as reducing noise, as well as reducing weight and cost due to the fewer fan blades.

More particularly, with the low solidity turbofan design described herein, the inventors have found a substantial reduction in flow blockage at the passage throats (described in more detail below), which offset the decreased solidity effect on aerodynamic performance. The inventors have further found that lower solidity through reduced blade number is beneficial to cruise efficiency, whereas lower solidity through reduction of the chord to diameter C/D ratio would be detrimental to cruise efficiency.

Furthermore, the inventors have found that maintaining the aerodynamic sweep near the airfoil hub or root at a relatively low level in the aft direction may be used for minimizing the required twist in the airfoil, which in turn moderates or reduces the stresses in the blade root and dovetail. This is particularly beneficial for blades of composite metal and carbon fiber construction.

Referring to FIG. 7, as discussed, adjacent airfoils 500 define circumferentially therebetween corresponding flow passages 552 for pressurizing the air 504 during operation. Each of the airfoils 500 includes stagger (or twist) represented by the stagger angle A from the axial or longitudinal axis.

The stagger increases from the root to the tip of the airfoil. For example, the stagger angle A at the blade tip may be relatively substantial (e.g., about 60 degrees) to position the leading edge 510 of one airfoil circumferentially adjacent but axially spaced from a suction side 522 of the next adjacent airfoil aft from the leading edge thereof to define a corresponding mouth 554 for the flow passage between the opposing pressure and suction sides 522, 524 of the adjacent airfoils. The contours and stagger of the adjacent airfoils over the radial span of the blades cause each flow passage to converge or decrease in flow area to a throat 556 of minimum flow area spaced aft from the mouth along most, if not all, of the radial span.

As further illustrated in FIG. 7, a relatively high airfoil stagger A also positions the trailing edge 512 of one airfoil 500 circumferentially adjacent to the pressure side 522 of the next adjacent airfoil while also being spaced axially therefrom in the tip region to define a corresponding discharge or outlet 558 for the corresponding flow passage between adjacent airfoils. In this way, the incoming air 504 is channeled in the corresponding flow passages 552 between adjacent airfoils as they rotate clockwise in FIG. 7 for pressurizing the air to produce the propulsion thrust during operation.

As shown in FIG. 7, a sloped platform 560 is disposed between corresponding pairs of adjacent airfoils 500 to conform with or match roots 502 of the blades, which may be in the form of sloped roots 502, that define the radially inner boundary of the fan air flowpath. Each platform 560 may be a discrete or separate component suitably mounted to the supporting fan disk between adjacent airfoils, or may be formed in halves integrally formed with the opposite sides of each fan blade along the roots thereof. The platform 560 slopes radially outwardly in the downstream or aft direction between the leading and trailing edges of the blade at a slope angle B (shown in FIG. 9) of about 15-20 degrees, and defines the radially inner boundary of each flow passage 552 between the blades. The sloped platform 560 cooperates with the airfoils themselves in pressurizing the airflow flowing downstream therepast.

The radially inner portion of an airfoils typically operates with subsonic airflow thereover for supercharging or initially pressurizing the inner portion of the air as it enters the booster compressor. The outer portion of the fan blade has the outer tip disposed at a suitably large outer diameter and rotates at a corresponding rotational speed for effecting supersonic airflow therepast during certain portions of the operating envelope of the aircraft being powered by the engine.

Another significant feature of the airfoil that affects its aerodynamic performance is its camber, which represents the amount of curvature of the airfoil along the radial or transverse sections between the leading and trailing edges thereof. FIG. 8 illustrates in solid line two graphs of exemplary camber and slope thereof over the span of the airfoil 500 from its root 502 at zero percent span to its tip 526 at 100 percent span, along with two representative radial sections of the airfoil at 70 percent and 100 percent span.

The airfoil camber may be defined by the difference in local axial inlet angle E at the leading edge 510 and the local axial exit angle F at the trailing edge 512. The camber line of the airfoil is the mean line extending between the leading and trailing edge of each section between the opposite pressure and suction sides. At the leading edge, the camber line defines the inlet angle E relative to the axial or longitudinal axis of the engine, and at the trailing edge, the camber line defines the exit angle F also relative to the axial axis.

The camber for each radial section of the airfoil may be simply obtained by subtracting the exit angle F from the inlet angle E, with the resulting camber represented in degrees as shown in the camber graph from root to tip of the airfoil. The graph illustrates that the camber for the fan blade 40 decreases between the root and tip of the airfoil in a substantially smooth curve from root to just short of the tip, with a slight local increase or peak in camber in the outer span of the airfoil immediately below the tip.

As indicated above, the fan blade 40 illustrated in FIG. 7 has substantial twist or stagger A of about 60 degrees for example from root to tip, along with a corresponding variation in camber of the individual radial sections from root to tip, and with a relatively large chord. The airfoil 500 of the fan blade is specifically configured to distinctly include sharp leading and trailing edges 510, 512 from root to tip which gradually increase in airfoil thickness to a maximum thickness along the midchord region between the opposite leading and trailing edges.

For example, the sharp leading edge 510 of the airfoil may be defined by a circumscribed square outline in the exemplary range of 26 to 42 mils (0.66-1.0 mm) thickness, with the sharp trailing edge 512 being represented by an inscribed circle of diameter 28 to 40 mils (0.7-1.0 mm). The maximum thickness of the airfoil correspondingly ranges from about 120 to 260 mils (3-6.6 mm) from root to tip, which maximum thickness varies along the chord in the exemplary range of 40 percent to 60 percent from the leading edge in axial projection of the twisted airfoil along the longitudinal engine axis.

As shown in FIG. 7, the stagger A and length of the chord C combine to effect the three-dimensional (3D) configuration of each airfoil. The section chords of the airfoil typically increase in length outboard from the root 502 to correspondingly barrel the airfoil above the root. The airfoil or chord barrelling may be observed in the axial side projection, which locally enlarges the midspan region of the airfoil preferably along both the leading and trailing edges 510, 512. The maximum airfoil barrel occurs at a suitable midspan region of the airfoil at an intermediate radial section of about 40 percent span from the root in the exemplary embodiment illustrated.

The leading edge barrelling extends in axial projection the leading edge upstream or forward of a straight line extending between the root and tip at the leading edge, and correspondingly the trailing edge barrelling extends in axial projection the trailing edge downstream or aft of a straight line extending between the root and tip at the trailing edge. In this way, the leading edge in the barrel extends axially forward of the airfoil root, and the trailing edge is correspondingly barrelled and also extends axially aft from the root.

The airfoil barrelling is yet another feature of the airfoil which affects the 3D configuration thereof and its aerodynamic performance in pressurizing the airflow channeled thereover during operation. In this regard, aerodynamic sweep is a conventional parameter represented by a local sweep angle which is a function of the direction of the incoming air and the orientation of the airfoil surface in both the axial and circumferential, or tangential, directions.

FIG. 9 illustrates the fan blade 40 along with the preferred aerodynamic sweep angle thereof as represented by the upper case letter S which has a negative value (−) for forward sweep, and a positive value (+) for aft sweep. A graph of aerodynamic sweep S for the leading edge 510 from root at zero percent span to tip at 100 percent span is superimposed over the airfoil.

The fan airfoil 500 preferably includes forward aerodynamic sweep S− at the airfoil tip 526 from leading edge 510 to trailing edge 512. Chord barrelling of the airfoil in conjunction with the forward tip sweep has significant aerodynamic benefits including increased flow capacity at high or maximum fan speed, while also improving part speed efficiency and stability margin.

However, the fan blade illustrated in FIG. 9 preferably includes non-forward aerodynamic sweep between the airfoil maximum barrel radial section and the sloped root 502 for further improving performance of the fan blade, particularly for part speed operability favoring cruise operation of the engine in propelling the aircraft in flight.

In the preferred embodiment illustrated in FIG. 9, the airfoil 500 includes relatively low aerodynamic sweep in the hub region near the root 502 or platform 560 which is less than about half the maximum sweep in the airfoil thereabove.

It is noted that the radial span of the airfoil is relatively large in the turbofan illustrated in FIG. 1, with the inner portion or panel of the airfoil below the midspan operating with relatively low Mach velocity of the airflow as it is pressurized or supercharged prior to entry in the booster compressor 20. The radially outer portion or panel of each airfoil above the midspan is highly twisted and is operated in transonic to supersonic Mach velocity of the airflow for providing a substantial pressure increase in the fan air used for generating the substantial amount of propulsion thrust from the engine.

The relatively low level of aerodynamic sweep which is preferably aft sweep near the airfoil root is found in the low Mach number region of the blade and serves to moderate the twist and leading edge stress in the blade root for mechanical benefits and to improve the pressurizing capability for a given level of camber.

FIG. 9 illustrates that the initially forward sweep S− at the airfoil tip 526 transitions through zero sweep to aft aerodynamic sweep S+ inwardly therefrom in the airfoil barrel both along the leading and trailing edges. In particular, the airfoil has maximum forward sweep along the leading edge at the tip 526 which transitions to maximum aft sweep inwardly therefrom along the leading edge in the barrel region of the airfoil. From the maximum aft sweep, the sweep decreases in aft magnitude along the leading edge towards the root 502 at zero span.

In the preferred embodiment illustrated in FIG. 9, the airfoil includes zero sweep along the leading edge 510 at the root 502, along with a local smaller peak in aft sweep along the leading edge between the root 502 and the maximum aft sweep. The maximum aft sweep along the leading edge occurs at about 60 percent span from the root, with the sweep remaining aft in magnitude over the entire inner panel span from 60 percent down to the root where it returns to zero sweep.

The leading edge sweep illustrated in FIG. 9 decreases rapidly from the maximum aft sweep near the midspan to the relatively low aft sweep in the inner panel down to the root which is substantially less than the magnitude of the maximum aft sweep. The local peak in leading edge sweep occurs at about five percent span and is less than half the magnitude of the maximum aft sweep, with the aft sweep being substantially lower in magnitude over most of the inner panel to about the 40 percent span elevation.

FIG. 8 illustrates that the camber decreases in general from root to tip in both parent and derived fan blades, with a similar decrease in the inner panel and less decrease over the outer panel of the derived blade relative to the parent blade. The rate or slope of camber decrease is therefore correspondingly different, and significantly locally increases camber in the outer panel between the midspan and tip.

As indicated above with respect to FIG. 8, the airfoil camber is defined by the difference in the inlet and exit angles (E-F). The inlet angle E preferably increases smoothly over the span of the airfoil from root to tip, with the exit angle F of the airfoil increasing from root to tip at a greater rate with about twice the overall magnitude. The increase in exit angle from root to tip may be used to correspondingly decrease the camber from the root toward the tip, with the camber graph in FIG. 8 illustrating a local increase in camber in the outer panel immediately below the airfoil tip down to about the midspan within the barrel.

FIG. 8 further illustrates that the camber has a greater rate of decrease, or slope, in the inner panel radially outboard from the root into the barrel to about 50 percent span for example than in the outer panel outboard from within the barrel at the 50 percent span for example toward the radially outer tip, which correspondingly locally increases the camber near or just below the airfoil tip over most of the outer panel.

The slope graph of camber illustrated in FIG. 8 shows the pronounced difference in a derived fan blade 40 over the preexisting blade 562 for introducing the local increase or peak in camber in the outer panel just below the airfoil tip 526. The slope represents the change in camber over the radial span of the airfoil from root to tip.

The slope in both curves is initially negative from the root outward as the camber decreases in magnitude outwardly along the span. The decreasing slope is similar in the two curves in the lower panel of the airfoils, and is affected in large part by the specific blade count in the full complement of fan blades in the fan row.

However, from the midspan region outwardly above the barrel in the outer panel, the camber and slope thereof in the solid and dashed curves change significantly. In the dashed curves, the camber decreases outwardly over the inner panel, and then increases locally over the outer panel. The slope or rate of camber change has a negative peak at 20% span in the inner panel, and transitions to a positive peak at the airfoil tip after passing through the zero value at about 73% span. This corresponds with a local minimum peak in camber for the preexisting fan blade.

In contrast, the slope of the camber in the derived fan blade over the preexisting blade, although generally similar over the inner panel, is remarkably different over the outer panel so that the camber has a local increase or peak in magnitude at about 80% span between the tip 526 and barrel.

The camber as shown by the solid curve varies in slope with a local negative magnitude or peak in the slope immediately below the tip 526 at about 94% span, followed by a local positive magnitude or peak in the slope at about 70% span therebelow. The solid slope curve therefore crosses the zero line at two locations above the midspan at about 62% and 74% span. And, below the midspan, the solid slope curve has a maximum negative peak at about 15% span.

Accordingly, the camber and its slope in the derived fan blade 40 are specifically configured to effect the local increase in camber in the outer panel just below the tip, with the camber level returning to its minimum value at the airfoil tip 526. In this way, a significant improvement in fan performance and efficiency may be obtained, while reducing efficiency losses at the airfoil tip itself which has a camber level closely similar to conventional practice.

As discussed, the fan blade with this special combination of aerodynamic sweep in its leading edge and locally increased camber in the outer panel is effective for increasing hub supercharging of the fan air while maintaining aeromechanical stability. The increased camber of the airfoil outer panel is further effective for delaying flutter and tailors the efficiency characteristics of the fan for improvement particularly at part speed corresponding with aircraft cruise operation. Maintaining the aerodynamic sweep near the airfoil hub or root at a relatively low level in the aft direction may be used for minimizing the required twist in the airfoil which in turn will moderate or reduce the stresses in the blade root and dovetail, which is particularly beneficial for blades of composite metal and carbon fiber construction.

As a consequence of the moderate aft sweep in the airfoil hub region, the tendency for radial outflow of the air being pressurized will be reduced for providing more turning in the air from a given camber level. The increased camber in the outer panel of the airfoil will correspondingly reduce the incidence angle of the incoming air and correspondingly improve aeromechanical characteristics in the part speed region of the operating envelope.

The additional camber of the outer panel of the airfoil may reduce high speed performance of the fan, but can be balanced by the improved performance at part speed, specifically cruise operation of the engine for maximizing overall efficiency of the fan.

Aerodynamic efficiency may be further improved in the turbofan engine 10 illustrated in FIG. 1 by reducing the solidity at the airfoil tips by reducing the number of blades from twenty-two to either twenty or eighteen, for example, while maintaining a similar or lager ratio of the tip chord over the tip diameter C/D in the derived fan 14 as originally found in the preexisting fan.

Furthermore, the reduction in number of fan blades increases the circumferential pitch P between the airfoils and increases the flow area of the flow passages 552, in particular at the throats 56 thereof, for reducing flow blockage during operation, and specifically at the airfoil tips subject to supersonic operation.

The configuration of the flow passage 552 illustrated in FIG. 7 is particularly important to efficient operation of the fan, and in particular at the airfoil tips subject to supersonic flow. The specific profiles of the pressure and suction sides of the individual airfoils, the lateral thickness T of the airfoil, the root to tip stagger and camber of the airfoils and, of course, the reduced solidity due to the reduction in blade count while maintaining equal the chord to diameter C/D ratio are all used to define each flow passage 552.

In particular, the airfoil tips 500 are locally angled and vary in thickness T or width between the leading and trailing edges 510, 512 to typically converge the flow passage 552 at the airfoil tips from the mouth 554 to the throat 556 and then diverge the flow passage also at the tip from the throat 556 to the outlet 558. Alternatively, the mouth and throat of the flow passages at the airfoil tips may be coincident in one plane at the leading edges, with the flow passages still diverging aft from the throats at the leading edges to the passage outlets at the trailing edges.

The convergence angle or slope between the mouth and the throat, and the divergence angle or slope between the throat and the outlet may be specifically designed for maximizing efficiency during supersonic operation of the blade tips in which aerodynamic shock is generated as the airflow is reduced in speed in the converging portion to choked flow of Mach 1 at the throat 556 followed in turn by subsonic diffusion in the diverging portion of the flow passage from or aft of the throat to the outlet.

The ratio of the flow area at the passage outlet 558 over the flow area at the throat 556 is a conventional measure of effective camber of the airfoils. The actual amount of airfoil camber near the tips thereof may be increased slightly over a conventional turbofan design as indicated above to allow the turbofan to tolerate the lower tip solidity during part-speed operation.

As indicated above, a modern turbofan is designed for an operating range from takeoff to cruise to landing, with cruise operation being predominant and for which maximum efficiency and operability is desired. However, part-speed performance must also be considered in good turbofan design and accommodated by the higher camber introduced near the blade tips for the low solidity turbofan design.

As discussed herein, part-speed operability may be improved by increasing the camber of the airfoils 500 at the tips 526 thereof in conjunction with the reduction in solidity by reduction in blade count.

As improved efficiency of the fan may be obtained through lowering solidity, the turbofan design may itself be otherwise conventional except as modified in accordance with the present disclosure. For example, the airfoils 500 described herein may be relatively large in diameter for supersonic tip operation in a modern turbofan engine with a substantial pressure ratio of about 1.5. The corresponding bypass ratio of the fan air which bypasses the core engine may be about 7.5 or greater. Further, the airfoils may be provided with suitable aerodynamic sweep which is preferably forward or negative (S−) at the tips 526 of the airfoils, with non-forward sweep near the hub or root.

This written description uses examples to disclose the present disclosure, including the best mode, and also to enable any person skilled in the art to practice the 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.

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

A turbomachine for an aircraft, comprising: an annular casing; and a fan disposed inside the annular casing and mounted for rotation about an axial centerline, the fan including fan blades that extend radially outwardly toward the annular casing; wherein the fan includes an average fan chord width of the fan blades (“c”), a diameter of the fan (“D”), a fan pressure ratio (“FPR”), and a redline corrected fan tip Mach number

( “ M tip , c ( RL ) ” ) ,

according to a First Performance Factor; wherein

FPF = [ c 0.15 · D ] ⁢ / [ [ FPR - 1 0.4 ] / M tip , c ( RL ) ] - 1.23 , and ⁢ wherein m 1 · [ M tip , c ( RL ) - 1.1 ] + 6 > FPF > m 1 · [ M tip , c ( RL ) - 1.1 ] + Δ ⁢ y 1 , and ⁢ wherein ⁢ 0 < Δ ⁢ y 1 < 6.

The turbomachine of one or more of these clauses wherein

M tip , c ( RL )

is within a range equal to or greater than 0.45 and equal to or less than 1.34.

The turbomachine of one or more of these clauses wherein

M tip , c ( RL )

is within a range equal to or greater than 0.45 and equal to or less than 1.12.

The turbomachine of one or more of these clauses wherein m₁ is equal to 0.87 when

M tip , c ( RL )

is greater than or equal to 1.1.

The turbomachine of one or more of these clauses wherein m₁ is equal to 1.0 when

M tip , c ( RL )

is greater than or equal to 1.1.

The turbomachine of one or more of these clauses wherein m₁ is equal to 2.5 when

M tip , c ( RL )

is less than 1.1.

The turbomachine of one or more of these clauses wherein m₁ is equal to 3.34 when

M tip , c ( RL )

is less than 1.1.

The turbomachine of one or more of these clauses wherein Δy₁ is equal to or greater than 0.0125 and less than 6.

The turbomachine of one or more of these clauses wherein Δy₁ is equal to or greater than 0.04 and less than 6.

The turbomachine of one or more of these clauses wherein Δy₁ is equal to or greater than 0.07 and less than 6.

The turbomachine of one or more of these clauses wherein Δy₁ is equal to or greater than 0.1 and less than 6.

The turbomachine of one or more of these clauses wherein Δy₁ is equal to or greater than 0.2 and less than 6.

The turbomachine of one or more of these clauses wherein the turbomachine has a gear ratio within a range equal to or greater than 3.2 and equal to or less than 5.0.

The turbomachine of one or more of these clauses wherein the fan blade includes an airfoil portion made from a polymer matrix composite (PMC) material, a metallic leading edge, the ratio c/D is between 0.16 and 0.21, and the fan has between 16 and 25 fan blades.

The turbomachine of one or more of these clauses wherein ratio c/D is within a range equal to or greater than 0.1 and equal to or less than 0.3.

The turbomachine of one or more of these clauses wherein ratio c/D is within a range equal to or greater than 0.15 and equal to or less than 0.21.

The turbomachine of one or more of these clauses wherein ratio c/D is equal to or greater than 0.1.

The turbomachine of one or more of these clauses wherein ratio c/D is equal to or less than 0.3.

The turbomachine of one or more of these clauses wherein FPR is within a range equal to or greater than 1.2 and equal to or less than 1.5.

The turbomachine of one or more of these clauses wherein FPR is within a range equal to or greater than 1.3 and equal to or less than 1.45.

The turbomachine of one or more of these clauses wherein FPR is equal to or greater than 1.2.

The turbomachine of one or more of these clauses wherein FPR is equal to or greater than 1.3.

The turbomachine of one or more of these clauses wherein FPR is equal to or less than 1.5.

The turbomachine of one or more of these clauses wherein FPR is equal to or less than 1.45.

The turbomachine of one or more of these clauses wherein U_c(tip) is within a range equal to or greater than 500 ft/sec and equal to or less than 1,500 ft/sec.

The turbomachine of one or more of these clauses wherein U_c(tip) is within a range equal to or greater than 500 ft/sec and equal to or less than 1,250 ft/sec.

A turbomachine comprising: an annular casing; and a fan disposed inside the annular casing and mounted for rotation about an axial centerline, the fan including fan blades that extend radially outwardly toward the annular casing; wherein the fan includes a fan hub-to-tip ratio (“HTR”), a fan blade count (“BC”), a fan pressure ratio (“FPR”), and a redline corrected redline fan tip Mach number

( “ M tip , c ( RL ) ” )

according to a Second Performance Factor (“SPF”),

SPF = π 4 ⁢ ( 1 - HTR 2 ) / ( BC 2 ⁢ 0 ) / ( FPR - 1 0.4 ) / M tip , c ( RL ) - 0.97 ; wherein m 2 · [ M tip , c ( RL ) - 1.1 ] + 1.5 > SPF > m 2 · [ M tip , c ( RL ) - 1.1 ] + Δ ⁢ y 2 , and ⁢ wherein 0 < Δ ⁢ y 2 < 1.5 .

The turbomachine of one or more of these clauses wherein Δy₂ is equal to or greater than 0.0075 and less than 1.5.

The turbomachine of one or more of these clauses wherein Δy₂ is equal to or greater than 0.01 and less than 1.5.

The turbomachine of one or more of these clauses wherein Δy₂ is equal to or greater than 0.02 and less than 1.5.

The turbomachine of one or more of these clauses wherein Δy₂ is equal to or greater than 0.024 and less than 1.5.

The turbomachine of one or more of these clauses wherein Δy₂ is equal to or greater than 0.037 and less than 1.5.

The turbomachine of one or more of these clauses wherein Δy₂ is equal to or greater than 0.04 and less than 1.5.

The turbomachine of one or more of these clauses wherein Δy₂ is equal to or greater than 0.06 and less than 1.5.

The turbomachine of one or more of these clauses wherein m₂ is equal to 0.41 when

M tip , c ( RL )

is greater than or equal to 1.1.

The turbomachine of one or more of these clauses wherein m₂ is equal to 0.55 when

M tip , c ( RL )

is less than 1.1.

The turbomachine of one or more of these clauses wherein HTR is within a range equal to or greater than 0.1 and equal to or less than 0.5.

The turbomachine of one or more of these clauses wherein HTR is within a range equal to or greater than 0.2 and equal to or less than 0.275.

The turbomachine of one or more of these clauses wherein HTR is within a range equal to or greater than 0.2 and equal to or less than 0.25.

The turbomachine of one or more of these clauses wherein HTR is equal to or greater than 0.1.

The turbomachine of one or more of these clauses wherein HTR is equal to or less than 0.5.

The turbomachine of one or more of these clauses wherein BC is within a range equal to or greater than 10 and equal to or less than 18.

The turbomachine of one or more of these clauses wherein FPR is within a range equal to or greater than 1.2 and equal to or less than 1.5.

The turbomachine of one or more of these clauses wherein FPR is within a range equal to or greater than 1.3 and equal to or less than 1.45.

A method of assembly, comprising: mounting a fan inside an annular casing for rotation about an axial centerline, the fan including fan blades that extend radially outwardly toward the annular casing; wherein the fan includes an average fan chord width of the fan blades (“c”), a diameter of the fan (“D”), a fan pressure ratio (“FPR”), and a redline corrected fan tip redline Mach number

( “ M tip , c ( RL ) ” )

according to a First Performance (“FPF”), wherein

FPF = [ c 0.15 · D ] / [ [ FPR - 1 0.4 ] / M tip , c ( RL ) ] - 1 . 2 ⁢ 3 ; m 1 · [ M tip , c ( RL ) - 1.1 ] + 6 > FPF > m 1 · [ M tip , c ( RL ) - 1.1 ] + Δ ⁢ y 1 ; and 0 < Δ ⁢ y 1 < 6 ;

or wherein the fan includes a fan hub-to-tip ratio (“HTR”), a fan blade count (“BC”), FPR, and

M tip , c ( RL )

according to a Second Performance Factor (“SPF”), wherein:

SPF = π 4 ⁢ ( 1 - HTR 2 ) / ( BC 2 ⁢ 0 ) / ( FPR - 1 0.4 ) / M tip , c ( RL ) - 0.97 ; m 2 · [ M tip , c ( RL ) - 1.1 ] + 1.5 > SPF > m 2 · [ M tip , c ( RL ) - 1.1 ] + Δ ⁢ y 2 ; and 0 < Δ ⁢ y 2 < 1.5 .

A turbomachine for an aircraft, comprising: an annular casing; and a fan disposed inside the annular casing and mounted for rotation about an axial centerline, the fan including fan blades that extend radially outwardly toward the annular casing; wherein the fan includes an average fan chord width of the fan blades (“c”), a diameter of the fan (“D”), a fan pressure ratio (“FPR”), and a redline corrected fan tip Mach number

( “ M tip , c ( RL ) ” ) ,

according to a First Performance Factor; wherein

PF = [ c 0.15 · D ] / [ [ FPR - 1 0.4 ] / M tip , c ( RL ) ] - 1.23 , and ⁢ wherein m 1 · [ M tip , c ( RL ) - 1.1 ] + 9 .14 > FPF > m 2 · [ M tip , c ( RL ) - 1.1 ] ,

wherein m₁ is equal to 9.43 when

M tip , c ( RL )

is greater than or equal to 1.1 and is equal to 27.02 when

M tip , c ( RL )

is less than 1.1, and wherein m₂ is equal to 0.87 when

M tip , c ( RL )

is greater than or equal to 1.1 and is equal to 3.34 when

M tip , c ( RL )

is less than 1.1.

The turbomachine of one or more of these clauses wherein FPF is within a range equal to or greater than −0.8 and equal to or less than 8.4.

The turbomachine of one or more of these clauses wherein FPF is within a range equal to or greater than 0 and equal to or less than 6.

The turbomachine of one or more of these clauses wherein FPF is within a range equal to or greater than 1 and equal to or less than 2.

The turbomachine of one or more of these clauses wherein

M tip , c ( RL )

is within a range equal to or greater than 0.8 and equal to or less than 1.5.

The turbomachine of one or more of these clauses wherein

M tip , c ( RL )

is within a range equal to or greater than 0.9 and equal to or less than 1.4.

The turbomachine of one or more of these clauses wherein ratio c/D is within a range equal to or greater than 0.1 and equal to or less than 0.3.

The turbomachine of one or more of these clauses wherein FPR is within a range equal to or greater than 1.2 and equal to or less than 1.6.

The turbomachine of one or more of these clauses wherein FPR is within a range equal to or greater than 1.3 and equal to or less than 1.5.

The turbomachine of one or more of these clauses wherein FPR is within a range equal to or greater than 1.35 and equal to or less than 1.45.

The turbomachine of one or more of these clauses wherein the turbomachine has a gear ratio within a range equal to or greater than 3.2 and equal to or less than 5.0.

The turbomachine of one or more of these clauses wherein the fan blade includes an airfoil portion made from a polymer matrix composite (PMC) material, a metallic leading edge, the ratio c/D is between 0.16 and 0.21, and the fan has between 16 and 25 fan blades.

A turbomachine comprising: an annular casing; and a fan disposed inside the annular casing and mounted for rotation about an axial centerline, the fan including fan blades that extend radially outwardly toward the annular casing; wherein the fan includes a fan hub-to-tip ratio (“HTR”), a fan blade count (“BC”), a fan pressure ratio (“FPR”), and a redline corrected redline fan tip Mach number

( “ M tip , c ( RL ) ” )

according to a Second Performance Factor (“SPF”),

SPF = SPF = π 4 ⁢ ( 1 - HTR 2 ) / ( BC 2 ⁢ 0 ) / ( FPR - 1 0.4 ) / M tip , c ( RL ) - 0.97 , wherein m 3 · [ M tip , c ( RL ) - 1.1 ] + 2.52 > SPF > m 4 · [ M tip , c ( RL ) - 1 . 1 ]

wherein m₃ is equal to 3.17, and wherein m₄ is equal to 0.41 when

M tip , c ( RL )

is greater than or equal to 1.1 and is equal to 0.55 when

M tip , c ( RL )

is less than 1.1.

The turbomachine of one or more of these clauses wherein SPF is within a range equal to or greater than 0.087 and equal to or less than 2.4.

The turbomachine of one or more of these clauses wherein SPF is within a range equal to or greater than 1 and equal to or less than 2.

A turbomachine for an aircraft, comprising: an annular casing; and a fan disposed inside the annular casing and mounted for rotation about an axial centerline, the fan including fan blades that extend radially outwardly toward the annular casing, a fan blade including: ****; wherein the fan includes an average fan chord width of the fan blades (“c”), a diameter of the fan (“D”), a fan pressure ratio (“FPR”), and a redline corrected fan tip Mach number

( “ M tip , c ( RL ) ” ) ,

according to a First Performance Factor; wherein

FPF = [ c 0.15 · D ] / [ [ ( FPR - 1 0.4 ] / M tip , c ( RL ) ] - 1.23 , and ⁢ wherein m 1 · [ M tip , c ( RL ) - 1.1 ] + 9.14 > FPF > m 2 · [ M tip , c ( RL ) - 1 .1 ] ,

wherein m₁ is equal to 9.43 when

M tip , c ( RL )

is greater than or equal to 1.1 and is equal to 27.02 when

M tip , c ( RL )

is less than 1.1, and wherein m₂ is equal to 0.87 when

M tip , c ( RL )

is greater than or equal to 1.1 and is equal to 3.34 when

M tip , c ( RL )

is less than 1.1.

The turbomachine of one or more of these clauses wherein

m 1 · [ M tip , c ( RL ) - 1.1 ] + 9.14 > FPF > m 2 · [ M tip , c ( RL ) - 1.1 ] + Δ ⁢ y 1 .

The turbomachine of one or more of these clauses wherein

m 3 · [ M tip , c ( RL ) - 1.1 ] + 2.52 > SPF > m 4 · [ M tip , c ( RL ) - 1.1 ] + Δ ⁢ y 2 .

A turbomachine for an aircraft, comprising: an annular casing; and a fan disposed inside the annular casing and mounted for rotation about an axial centerline, the fan including fan blades that extend radially outwardly toward the annular casing, a fan blade including: an airfoil extending outwardly from a sloped root and including opposite pressure and suction sides extending longitudinally in span from the sloped root to an opposite tip, and extending axially in chord between opposite leading and trailing edges, the airfoil further including stagger increasing between the sloped root and tip, camber decreasing therebetween, and chord length increasing outboard from the sloped root to barrel the airfoil along both the leading and trailing edges, the airfoil further including maximum forward aerodynamic sweep at the tip, and non-forward aerodynamic sweep between a maximum barrel in the airfoil and the sloped root; wherein the fan includes an average fan chord width of the fan blades (“c”), a diameter of the fan (“D”), a fan pressure ratio (“FPR”), and a redline corrected fan tip Mach number

( “ M tip , c ( RL ) ” ) ,

according to a First Performance Factor; wherein

FPF = [ c 0.15 · D ] ⁢ / [ [ FPR - 1 0.4 ] / M tip , c ( RL ) ] - 1.23 , and ⁢ wherein m 1 · [ M tip , c ( RL ) - 1.1 ] + 9.14 > FPF > m 2 · [ M tip , c ( RL ) - 1.1 ] ,

wherein m₁ is equal to 9.43 when

M tip , c ( RL )

is greater than or equal to 1.1 and is equal to 27.02 when

M tip , c ( RL )

is less tian 1.1, and wherein m₂ is equal to 0.87 when

M tip , c ( RL )

is greater than or equal to 1.1 and is equal to 3.34 when

M tip , c ( RL )

is less than 1.1.

A turbomachine comprising: an annular casing; and a fan disposed inside the annular casing and mounted for rotation about an axial centerline, the fan including fan blades that extend radially outwardly toward the annular casing, a fan blade including: an airfoil extending outwardly from a sloped root and including opposite pressure and suction sides extending longitudinally in span from the sloped root to an opposite tip, and extending axially in chord between opposite leading and trailing edges, the airfoil further including stagger increasing between the sloped root and tip, camber decreasing therebetween, and chord length increasing outboard from the sloped root to barrel the airfoil along both the leading and trailing edges, the airfoil further including maximum forward aerodynamic sweep at the tip, and non-forward aerodynamic sweep between a maximum barrel in the airfoil and the sloped root; wherein the fan includes a fan hub-to-tip ratio (“HTR”), a fan blade count (“BC”), a fan pressure ratio (“FPR”), and a redline corrected redline fan tip Mach number

( “ M tip , c ( RL ) ” )

according to a Second Performance Factor (“SPF”),

SPF = π 4 ⁢ ( 1 - HTR 2 ) / ( BC 20 ) / ( FPR - 1 0.4 ) / M tip , c ( RL ) - 0.97 , wherein m 3 · [ M tip , c ( RL ) - 1.1 ] + 2.52 > SPF > m 4 · [ M tip , c ( RL ) - 1.1 ]

wherein m₃ is equal to 3.17, and wherein m₄ is equal to 0.41 when

M tip , c ( RL )

is greater than or equal to 1.1 and is equal to 0.55 when

M tip , c ( RL )

is less than 1.1.

A turbomachine for an aircraft, comprising: an annular casing; and a fan disposed inside the annular casing and mounted for rotation about an axial centerline, the fan including fan blades that extend radially outwardly toward the annular casing, a fan blade including: an airfoil extending outwardly from a sloped root and including opposite pressure and suction sides extending longitudinally in span from the sloped root to an opposite tip, and extending axially in chord between opposite leading and trailing edges, the airfoil further including stagger increasing between the sloped root and tip, camber decreasing therebetween, and chord length increasing outboard from the sloped root to barrel the airfoil along both the leading and trailing edges, the airfoil further including maximum forward aerodynamic sweep at the tip, and non-forward aerodynamic sweep between a maximum barrel in the airfoil and the sloped root; wherein the fan includes an average fan chord width of the fan blades (“c”), a diameter of the fan (“D”), a fan pressure ratio (“FPR”), and a redline corrected fan tip Mach number

( “ M tip , c ( RL ) ” ) ,

according to a First Performance Factor; wherein

FPF = [ c 0.15 · D ] ⁢ / [ [ FPR - 1 0.4 ] / M tip , c ( RL ) ] - 1.23 , and ⁢ wherein m 1 · [ M tip , c ( RL ) - 1.1 ] + 9.14 > FPF > m 2 · [ M tip , c ( RL ) - 1.1 ] ,

wherein m₁ is equal to 9.43 when

M tip , c ( RL )

is greater than or equal to 1.1 and is equal to 27.02 when

M tip , c ( R ⁢ L )

is less than 1.1, and wherein m₂ is equal to 0.87 when

M tip , c ( RL )

is greater than or equal to 1.1 and is equal to 3.34 when

M tip , c ( RL )

is less than 1.1; wherein the fan includes a fan hub-to-tip ratio (“HTR”), a fan blade count (“BC”), a fan pressure ratio (“FPR”), and a redline corrected redline fan tip Mach number

( “ M tip , c ( R ⁢ L ) ” )

according to a Second Performance Factor (“SPF”),

SPF = π 4 ⁢ ( 1 - HTR 2 ) / ( B ⁢ C 2 ⁢ 0 ) / ( FPR - 1 0.4 ) / M tip , c ( RL ) - 0.97 , wherein ⁢ m 3 · [ M tip , c ( R ⁢ L ) - 1 . 1 ] + 2 .52 > SPF > m 4 · [ M tip , c ( R ⁢ L ) - 1 .1 ]

wherein m₃ is equal to 3.17, and wherein m₄ is equal to 0.41 when

M tip , c ( RL )

is greater than or equal to 1.1 and is equal to 0.55 when

M tip , c ( RL )

is less than 1.1.

The turbomachine of one or more of these clauses wherein the forward sweep at the tip transitions to a maximum aft sweep inwardly along the leading edge in the barrel, and decreases in aft magnitude to non-forward sweep at the root.

The turbomachine of one or more of these clauses wherein the airfoil further includes sharp leading and trailing edges from the root to tip and gradually increases to a maximum thickness along the midchord region therebetween.

The turbomachine of one or more of these clauses wherein the airfoil further includes low sweep near the root being less than about half the maximum sweep in the airfoil thereabove.

The turbomachine of one or more of these clauses wherein the forward sweep at the airfoil tip transitions to aft sweep inwardly therefrom in the airfoil barrel.

The turbomachine of one or more of these clauses wherein the airfoil further includes maximum forward sweep at the tip transitioning to maximum aft sweep inwardly along the leading edge in the barrel, and decreasing in aft magnitude toward the root.

The turbomachine of one or more of these clauses wherein the airfoil camber is defined by the difference in axial inlet angle at the leading edge and axial exit angle at the trailing edge, and the exit angle increases from root to tip to correspondingly decrease the camber from the root toward the tip.

The turbomachine of one or more of these clauses wherein the camber has a greater rate of decrease outboard from the root into the barrel than outboard the barrel toward the tip for locally increasing the camber near the tip.

The turbomachine of one or more of these clauses wherein the camber varies in slope with a local negative peak immediately below the tip followed by a local positive peak therebelow.

The turbomachine of one or more of these clauses wherein the camber has a local increase between the tip and barrel.

The turbomachine of one or more of these clauses wherein the airfoil includes zero sweep along the leading edge at the root.

The turbomachine of one or more of these clauses wherein the airfoil includes a local peak in aft sweep along the leading edge between the root and the maximum aft sweep.

A plurality of fan blades according to one or more of these clauses arranged circumferentially in a row to define corresponding flow passages between adjacent airfoils for pressurizing air; each of the airfoils including stagger increasing between the root and the tip to position the leading edge of one airfoil circumferentially adjacent to the suction side of the next adjacent airfoil to define a mouth for the flow passage therebetween, with the flow passage converging to a throat aft from the mouth; and adjacent airfoils including a sloped platform conforming to the sloped root for effecting with the airfoils the non-forward sweep near the platform.

A fan blade row according to one or more of these clauses wherein the airfoil stagger positions the trailing edge of one airfoil circumferentially adjacent to the pressure side of the next adjacent airfoil to define an outlet for the flow passage therebetween; and the airfoil tips vary in thickness between the leading and trailing edges to diverge the flow passage therebetween.

A fan blade row according to one or more of these clauses comprising no more than twenty of the fan blades having a solidity defined by the ratio of the airfoil chord over the circumferential pitch and being low in magnitude at the tips to position the leading edge of each tip circumferentially near the trailing edge of the next adjacent tip and correspondingly increase the width of the throat.

A method of improving aerodynamic efficiency of a fan blade according to one or more of these clauses comprising: deriving the blade from a preexisting fan blade; increasing the airfoil camber locally near the airfoil tip; and eliminating forward sweep along the leading edge near the airfoil root.