TURBOMACHINE AND METHOD OF ASSEMBLY

Abstract

A turbomachine includes an annular casing and a fan disposed inside the annular casing and mounted for rotation about an axial centerline. The annular casing is formed at least partially of a composite material. The fan includes fan blades that extend radially outwardly toward the annular casing. The fan has an average chord fan width according to a first performance factor. The fan has a quantity of fan blades according to a second performance factor.

Description

FIELD

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

BACKGROUND

In one form, a gas turbine engine may 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 gas turbine 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 gas turbine 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 gas turbine 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 gas turbine 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 schematic cross-sectional illustration of a gas turbine engine including a composite fan composite casing and inner shell with circumferential varying thickness;

FIG. 8 is an enlarged cross-sectional illustration of back sheet and face sheet used for the circumferential varying thickness of the composite inner shell illustrated in FIG. 7;

FIG. 9 is a schematic side view illustration of the back sheet and the face sheet in the circumferential varying thickness composite shell illustrated in FIG. 8;

FIG. 10 is a schematic perspective view illustration of the composite shell illustrated in FIG. 9;

FIG. 11 is a schematic side view illustration of the composite shell illustrated in FIG. 10 wrapped in Kevlar;

FIG. 12 is a schematic side view illustration of the composite shell without the Kevlar wrap illustrated in FIG. 11;

FIG. 13 is a schematic axial cross-sectional view through 13-13 in FIG. 11;

FIG. 14 is a schematic axial cross-sectional view through 14-14 in FIG. 11; and

FIG. 15 is a schematic axial cross-sectional view through 15-15 in FIG. 11.

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 may be a combination of at least two or more non-metallic elements or materials. Examples of a composite material may 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, may include several layers or plies of composite material. The layers or plies may 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 may be used in forming or coupling composite components. Adhesives may include resin and phenolics, wherein the adhesive may 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 may 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 may be generally classified as thermosets or thermoplastics. Thermoplastic resins are generally categorized as polymers that may 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 may include, but is not limited to, dry carbon fibers woven together with thermoplastic polymer fibers or filaments. Non-prepreg braided architectures may 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 may 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 may 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) may 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 may include prepreg, braided material, woven material, or any combination thereof.

Resin may 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 may require post-curing processing.

It is contemplated that RTM may be a vacuum assisted process. That is, the air from the cavity or mold may 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 may be manual or automated.

The dry fibers or matrix material may 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 may 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 may 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, such as a gas turbine engine 10, in accordance with an exemplary embodiment of the present disclosure. More particularly, for the embodiment of FIG. 1, the gas turbine engine 10 is a high-bypass turbofan jet engine. gas turbine engine 10 As shown in FIG. 1, the gas turbine 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 gas turbine 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 32. A high pressure (HP) shaft or spool 34 drivingly connects the HP turbine 28 to the HP compressor 24. A low pressure (LP) shaft or spool 36 drivingly connects the LP turbine 30 to the LP 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 the 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 actuator 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 actuator 44 is not present. The fan blades 40, disk 42, and actuator 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 may 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 may comprise a first rotational speed and the output may 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 may 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 may 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 may 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 may 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, such as an 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 gas turbine engine 10, a volume of air 58 enters the gas turbine 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 LP 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 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 HP turbine 28 and the LP 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 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 gas turbine engine 10, also providing propulsive thrust.

It should be appreciated, however, that the gas turbine 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 gas turbine 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 gas turbine engine 10 may include or be operably connected to any other suitable accessory systems. Additionally, or alternatively, the exemplary gas turbine 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 gas turbine 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 may 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 a high-pressure surface 112 opposite the low-pressure surface 110 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 local chord 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 lengths of the local chord 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 may 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 may 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:

$\begin{array}{cc}\mathrm{FPF}=\left[\frac{c}{0.15·D}\right]/\left[\left[\frac{\sqrt{FPR-1}}{\sqrt{0.4}}\right]/M_{\mathrm{tip},c}^{\left(RL\right)}\right]-1\text{.23}& \left(1\right)\end{array}$ $\begin{array}{cc}{m}_1·\left[M_{\mathrm{tip},c}^{\left(RL\right)}-1.1\right]+6>\mathrm{FPF}>m_1·\left[M_{\mathrm{tip},c}^{\left(RL\right)}-1.1\right]+\Delta y_1& \left(2\right)\end{array}$

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 “m₁·[M_tip,c^(RL)−1.1]+Δy₁” 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 may be 0.0125, 0.04, 0.07, 0.1, or 0.2, or may 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 gas turbine engine embodiments, of which gas turbine engines 210, 212, 214, 216 may be referred to as low speed engine designs (as indicated by subplot area 218), and gas turbine engines 220, 222, 224, 226 may be referred to as high speed engine designs (as indicated by subplot area 228). Each of the gas turbine 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, gas turbine engines 214, 216, 224, and 226 may provide advantages over gas turbine 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 C_L 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 m₁·[M_tip,c^(RL)−1.1]+0.0125, greater than m₁·[M_tip,c^(RL)−1.1]+0.04, greater than m₁·[M_tip,c^(RL)−1.1]+0.07, greater than m₁·[M_tip,c^(RL)−1.1]+0.1, or greater than m₁·[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:

$\begin{array}{cc}\mathrm{SPF}=\sqrt{\frac{\pi }{4}\left(1-HT{R}^2\right)/\left(\frac{BC}{20}\right)}/\left(\frac{\sqrt{FPR-1}}{\sqrt{0.4}}\right)/M_{t\mathrm{ip},c}^{\left(RL\right)}-0.97& \left(3\right)\end{array}$ $\begin{array}{cc}{m}_2·\left[M_{t\mathrm{ip},c}^{\left(RL\right)}-1.1\right]+1.5>\mathrm{SPF}>m_2·\left[M_{t\mathrm{ip},c}^{\left(RL\right)}-1.1\right]+\Delta y_2& \left(4\right)\end{array}$

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 fan pressure ratio and a redline corrected fan tip Mach tip,c 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 “m₂·[M_tip,c^(RL)−1.1]+Δy₂” 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 may be 0.0075, 0.01, 0.02, 0.024, 0.037, 0.04, or 0.06, or may 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 gas turbine engine embodiments, of which gas turbine engines 310, 312, 314, 316 may be referred to as low speed engine designs (as indicated by subplot area 318), and gas turbine engines 320, 322, 324, 326 may be referred to as high speed engine designs (as indicated by subplot area 328). Each of the gas turbine 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, gas turbine engines 314, 316, 324, and 326 may provide advantages over gas turbine 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 (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 SPF values greater than m₂·[M_tip,c^(RL)−1.1]+0.0075, greater than m₂·[M_tip,c^(RL)−1.1]+0.01, greater than m₂·[M_tip,c^(RL)−1.1]+0.02, greater than m₂·[M_tip,c^(RL)−1.1]+0.024, greater than m₂·[M_tip,c^(RL)−1.1]+0.037, greater than m₂·[M_tip,c^(RL)−1.1]+0.04, or greater than m₂·[M_tip,c^(RL)−1.1]+0.06 (these other examples are schematically represented by the phantom line 302).

FIG. 5 shows additional example gas turbine 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 “m₁” 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:

$\begin{array}{cc}\mathrm{FPF}=\left[\frac{c}{0.15·D}\right]/\left[\left[\frac{\sqrt{FPR-1}}{\sqrt{0.4}}\right]/M_{\mathrm{tip},c}^{\left(RL\right)}\right]-1\text{.23}& \left(5\right)\end{array}$ $\begin{array}{cc}{m}_1·\left[M_{\mathrm{tip},c}^{\left(RL\right)}-1.1\right]+9.14>\mathrm{FPF}>m_2·\left[M_{\mathrm{tip},c}^{\left(RL\right)}-1\text{.1}\right]& \left(6\right)\end{array}$

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:

$\begin{array}{cc}\mathrm{SPF}=\sqrt{\frac{\pi }{4}\left(1-HT{R}^2\right)/\left(\frac{BC}{20}\right)}/\left(\frac{\sqrt{FPR-1}}{\sqrt{0.4}}\right)/M_{t\mathrm{ip},c}^{\left(RL\right)}-0.97& \left(7\right)\end{array}$ $\begin{array}{cc}{m}_3·\left[M_{t\mathrm{ip},c}^{\left(RL\right)}-1.1\right]+2.52>\mathrm{SPF}>m_4·\left[M_{t\mathrm{ip},c}^{\left(RL\right)}-1\text{.1}\right]& \left(8\right)\end{array}$

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.

	TABLE 1
	Example	HTR	FPR	Mtip, c(RL)	SPF	FPF
	1	0.206	1.522	1.417	1.782	2.374
	2	0.400	1.376	1.421	0.981	0.976
	3	0.260	1.204	1.177	0.823	2.722
	4	0.224	1.595	0.976	0.646	−0.359
	5	0.213	1.517	0.815	0.613	−0.823
	6	0.265	1.448	1.497	1.152	1.161
	7	0.352	1.250	0.962	0.087	−0.445
	8	0.394	1.328	1.228	2.403	6.606
	9	0.213	1.517	0.815	0.613	−0.823
	10	0.235	1.240	1.231	2.053	8.398

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 will be appreciated from the description above, fan pressure ratios, corrected fan tip speeds, fan diameters, and hub-to-tip ratios related using the discovered relationships may be used to select a fan chord width, a blade count, or both to provide a gas turbine engine having improved engine aerodynamic efficiency and/or improved fuel efficiency. During the course of designing a more efficient gas turbine engine, the inventors found that providing a composite fan case for surrounding and circumscribing the fan blades provides additional advantages synergistically when utilizing the above relationships, including improving a durability of the gas turbine engine.

In additional embodiments, a fan section of the gas turbine engine includes a composite fan case that may prevent damage caused by impact of foreign objects, such as birds or debris. The fan case may also prevent crack formation caused by strains induced during large, applied loads. More specifically, the composite fan case may prevent consequential damage resulting from a blade fragment or an entire fan blades being dislodged by impact with foreign objects, such as during a blade release event, which may damage the surrounding fan case and cause circumferential cracking based on the resulting load on the fan.

The composite fan case disclosed herein may be integrated into the fan module contemplated by the above relationships (1) through (4) to further improve performance of the gas turbine engine. For example, the First Performance Factor (FPF) and Second Performance Factor (SPF) parameters drive the fan blade design towards a lower corrected fan tip speed (U_c(tip)), which in turn allows for the selection of a lower fan chord and potentially a reduced blade count (BC). This reduction in fan chord or tip speed, and particularly a reduction in both fan chord and tip speed, reduces kinetic energy during a blade release event and enables the use of a lighter fan case, which synergistically improves the efficiency of the engine by reducing the overall system weight and enhancing fuel efficiency. Moreover, the composite material forming the composite fan case may have a lower Coefficient of Thermal Expansion (CTE) than metallic materials. Accordingly, the composite material will expand less with increased temperatures than metallic materials, maintaining more desirable tip clearances with the fan blades without requiring a dedicated clearance control system.

Accordingly, the inventors of the present disclosure found that providing a composite fan case in combination with the above relationships (1) through (4) results in a gas turbine engine having synergistic engine improvement in terms of durability and fuel efficiency.

Illustrated schematically in FIG. 7 is an exemplary gas turbine engine 510 circumscribed about a centerline axis 508. The gas turbine engine 510 may be similar or analogous to the gas turbine engine 10 discussed above with respect to FIG. 1. For example, the gas turbine engine 510 includes, in downstream flow relationship, a fan 512 which receives ambient air 514, a low pressure booster or compressor 516, a high pressure compressor (HPC) 518, a combustor 520 which mixes fuel with the ambient air 514 pressurized by the HPC 518 for generating combustion gases 522 which flow downstream through a high pressure turbine (HPT) 524, and a low pressure turbine (LPT) 526 from which the combustion gases 522 are discharged from the gas turbine engine 510. A first or high pressure shaft 528 joins the HPT 524 to the HPC 518, and a second or low pressure shaft 530 joins the LPT 526 to both the fan 512 and the low pressure booster or compressor 516.

A fan section 546 of the gas turbine engine 510 includes the fan 512 and a fan case assembly 541 with a composite fan case 542 circumscribing and surrounding fan blades 544 of the fan 512. The fan case assembly 541 further included a metallic fan casing 545 aft or downstream of and bolted to the composite fan case 542. A fan blade containment system 540 circumscribes and surrounds the fan 512 and the fan blades 544 to retain any fan blades 544 or fan blade fragments dislodged from the fan 512. A “blade-out event” arises when a fan blade or portion thereof is accidentally released from a rotor of a high-bypass turbofan engine. When suddenly released during flight, a fan blade may impact a surrounding fan case with substantial force, and resulting loads on the fan case may cause circumferential cracking of the fan case.

FIGS. 7-9 illustrate the fan blade containment system 540 includes the composite fan case 542 circumscribing the fan blades 544. The fan blade containment system 540 includes a circumferentially varying thickness T portion 548 of the composite fan case 542. The composite material of the fan case 542 is a lightweight and high-strength material. The circumferentially varying thickness T around the composite fan case 542 is designed to guide strains induced during large, applied loads, such as, during fan blade-out (FBO) events. It is highly beneficial during an FBO event to drive subsequent case damage progression to areas of intended reinforcement such as a metallic fan casing flange 543 of the metallic fan casing 545. An exemplary circumferentially varying thickness T is illustrated by first and second thicknesses T1 and T2 at first and second circumferential positions P1, P2, respectively. It should be understood that, in some example embodiments, the fan blade containment system 540 including the composite fan case 542 may be incorporated into the gas turbine engine 10 described above with respect to FIG. 1.

Referring to FIGS. 9-12, the composite fan case 542 is annular and includes an annular composite shell 550 extending from a forward flange 552 aft or downstream to an aft flange 554. The composite shell's aft flange 554 is bolted to the metallic fan casing flange 543 of the metallic fan casing 545. An annular composite back sheet 556 is spaced radially outwardly of the annular composite shell 550 and surrounds the fan blades 544. An annular filler layer 560, such as honeycomb, is disposed between the annular composite shell 550 and the composite back sheet 556. The annular filler layer 560 is axially trapped between annular forward and aft seams 564, 566 between the annular composite shell 550 and composite back sheet 556. As illustrated in FIGS. 10, 11, 13, and 14, an annular layer of Kevlar 574 may cover and surround the annular composite back sheet 556 surrounding the fan blades 544.

The aft seams 566 includes aftwardly or downstream extending lobes or fingers 570 of the composite back sheet 556. The annular composite shell 550 contacts and is bonded to the composite back sheet 556, as illustrated in FIGS. 13 and 14, along the forward and aft seams 564, 566. This provides the circumferentially varying thickness T portion 548 of the composite fan case 542 in which the thicker first thickness T1 is that of the combined annular composite shell 550 and composite back sheet 556. The thinner second thickness T2 is that of just the composite shell 550 circumferentially between the lobes or fingers 570 of the composite back sheet 556 in the circumferentially varying thickness T portion 548 as illustrated in FIG. 15.

The circumferentially varying thickness T portion 548 of the composite fan case 542 may arrest circumferential cracks in the composite portion of the case and subsequent damage may be directed to metallic regions, outside of the primary composite containment region, where metallic reinforcement is available to arrest the crack and prevent circumferential composite case failure. The containment shell is designed to withstand damage during FBO, this design helps insure that damage does not propagate to an extended 360 degree crack.

The circumferentially varying thickness T of the composite shell makes for a torturous path for any crack propagation to happen. Thus, any propagation may be ‘steered’ axially to a region that has added reinforcements such as the metallic fan casing flange 543 of the metallic fan casing 545.

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 fan blade containment system comprising: a composite fan case, the composite fan case including an annular composite shell extending from a forward flange aft or downstream to an aft flange, and a circumferentially varying thickness portion of the composite fan case including the composite shell.

The system a of one or more of these clauses, further comprising annular forward and aft seams between the annular composite shell and a composite back sheet.

The system of one or more of these clauses, further comprising the circumferentially varying thickness portion of the composite fan case including aftwardly or downstream extending lobes or fingers of the composite back sheet on the composite shell in the aft seam.

The system of one or more of these clauses, further comprising the circumferentially varying thickness portion including a first thickness of the lobes or fingers thicker than a second thickness of the composite shell circumferentially between the lobes or fingers.

The system of one or more of these clauses, further comprising an annular composite back sheet spaced radially outwardly of the annular composite shell and an annular filler layer disposed radially therebetween.

The system of one or more of these clauses, further comprising the filler layer including honeycomb.

The system of one or more of these clauses, further comprising annular forward and aft seams between the annular composite shell and composite back sheet axially trapping the filler layer between the forward and aft seams.

The system of one or more of these clauses, further comprising the circumferentially varying thickness portion of the composite fan case including aftwardly or downstream extending lobes or fingers of the composite back sheet on the composite shell in the aft seam.

The system of one or more of these clauses, further comprising the circumferentially varying thickness portion including a first thickness of the lobes or fingers thicker than a second thickness of the composite shell circumferentially between the lobes or fingers.

A fan case assembly comprising: a metallic fan casing aft or downstream of and bolted to a composite fan case, the composite fan case including an annular composite shell extending from a forward flange aft or downstream to an aft flange, the metallic fan casing including a metallic fan casing flange bolted to the aft flange, and a circumferentially varying thickness portion of the composite fan case including the composite shell.

The fan case assembly of one or more of these clauses, further comprising an annular composite back sheet spaced radially outwardly of the annular composite shell and an annular filler layer disposed radially therebetween.

The fan case assembly of one or more of these clauses, further comprising the filler layer including honeycomb.

The fan case assembly of one or more of these clauses, further comprising annular forward and aft seams between the annular composite shell and composite back sheet axially trapping the filler layer between the forward and aft seams.

The fan case assembly of one or more of these clauses, further comprising the circumferentially varying thickness portion of the composite fan case including aftwardly or downstream extending lobes or fingers of the composite back sheet on the composite shell in the aft seam and a first thickness of the lobes or fingers thicker than a second thickness of the composite shell circumferentially between the lobes or fingers.

A gas turbine engine fan section comprising: a fan section including a fan, a fan case assembly surrounding the fan and including a metallic fan casing aft or downstream of and bolted to a composite fan case, a fan blade containment system including the composite fan case circumscribing and surrounding the fan and the fan blades, the composite fan case including an annular composite shell extending from a forward flange aft or downstream to an aft flange, the metallic fan casing including a metallic fan casing flange bolted to the aft flange, and a circumferentially varying thickness portion of the composite fan case including the composite shell.

The gas turbine engine fan section of one or more of these clauses, further comprising an annular composite back sheet spaced radially outwardly of the annular composite shell and an annular filler layer disposed radially therebetween.

The gas turbine engine fan section of one or more of these clauses, further comprising the filler layer including honeycomb.

The gas turbine engine fan section of one or more of these clauses, further comprising annular forward and aft seams between the annular composite shell and composite back sheet axially trapping the filler layer between the forward and aft seams.

The gas turbine engine fan section of one or more of these clauses, further comprising the circumferentially varying thickness portion of the composite fan case including aftwardly or downstream extending lobes or fingers of the composite back sheet on the composite shell in the aft seam and a first thickness of the lobes or fingers thicker than a second thickness of the composite shell circumferentially between the lobes or fingers.

The gas turbine engine fan section of one or more of these clauses, further comprising an annular layer of Kevlar covering the annular composite back sheet covering and surrounding the composite back sheet.

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 speed (“U_c(tip)”), according to a First Performance Factor; wherein FPF=[c/D]/[√{square root over (FPR−1)}/U_c(tip)]; and wherein 0.24*U_c(tip)+489>FPF>0.24*U_c(tip)−12+dy₁ and wherein 0<dy₁<500.

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.

The turbomachine of one or more of these clauses wherein dy₁ is equal to or greater than 7.5 and equal to or less than 500.

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

$\mathrm{FPF}=\left[\frac{c}{0.15·D}\right]/\left[\left[\frac{\sqrt{FPR-1}}{\sqrt{0.4}}\right]/M_{\mathrm{tip},c}^{\left(RL\right)}\right]-1.23,$

and wherein m₁·[M_tip,c^(RL)−1.1]+6>FPF>m₁·[M_tip,c^(RL)−1.1]+Δy₁, and wherein 0<Δy₁<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”),

$\mathrm{SPF}=\sqrt{\frac{\pi }{4}\left(1-HT{R}^2\right)}/\left(\frac{BC}{20}\right)/\left(\frac{\sqrt{FPR-1}}{\sqrt{0.4}}\right)/M_{t\mathrm{ip},c}^{\left(RL\right)}-0.97,$ $\mathrm{wherein}{m}_2·\left[M_{t\mathrm{ip},c}^{\left(RL\right)}-1.1\right]+1.5>\mathrm{SPF}>m_2·\left[M_{t\mathrm{ip},c}^{\left(RL\right)}-1.1\right]+\Delta y_2,$

and wherein 0<Δy₂<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 Factor (“FPF”), wherein:

$\mathrm{FPF}=\left[\frac{c}{0.15·D}\right]/\left[\left[\frac{\sqrt{\mathrm{FPR}-1}}{\sqrt{0.4}}\right]/M_{\mathrm{tip},c}^{\left(\mathrm{RL}\right)}\right]-1.23;$ $m_1·\left[M_{\mathrm{tip},c}^{\left(\mathrm{RL}\right)}-1.1\right]+6>\mathrm{FPF}>m_1·\left[M_{\mathrm{tip},c}^{\left(\mathrm{RL}\right)}-1.1\right]+\Delta y_1;$ $\mathrm{and}$ $0<\Delta y_1<6;$

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:

$\mathrm{SPF}=\sqrt{\frac{\pi }{4}\left(1-{\mathrm{HTR}}^2\right)/\left(\frac{\mathrm{BC}}{20}\right)}/\left(\frac{\sqrt{\mathrm{FPR}-1}}{\sqrt{0.4}}\right)/M_{\mathrm{tip},c}^{\left(\mathrm{RL}\right)}-0.97;$ $m_2·\left[M_{\mathrm{tip},c}^{\left(\mathrm{RL}\right)}-1.1\right]+1.5>\mathrm{SPF}>m_2·\left[M_{\mathrm{tip},c}^{\left(\mathrm{RL}\right)}-1.1\right]+\Delta y_2;$ $\mathrm{and}$ $0<\Delta y_2<1.5.$

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

$\mathrm{FPF}=\left[\frac{c}{0.15·D}\right]/\left[\left[\frac{\sqrt{\mathrm{FPR}-1}}{\sqrt{0.4}}\right]/M_{\mathrm{tip},c}^{\left(\mathrm{RL}\right)}\right]-1.23,$

and wherein m₁·[M_tip,c^(RL)−1.1]+9.14>FPF>m₂·[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”),

$\mathrm{SPF}=\mathrm{SPF}=\sqrt{\frac{\pi }{4}\left(1-{\mathrm{HTR}}^2\right)/\left(\frac{\mathrm{BC}}{20}\right)}/\left(\frac{\sqrt{\mathrm{FPR}-1}}{\sqrt{0.4}}\right)/M_{\mathrm{tip},c}^{\left(\mathrm{RL}\right)}-0.97,$ $\mathrm{wherein}$ $m_3·\left[M_{\mathrm{tip},c}^{\left(\mathrm{RL}\right)}-1.1\right]+2.52>\mathrm{SPF}>m_4·\left[M_{\mathrm{tip},c}^{\left(\mathrm{RL}\right)}-1.1\right]$

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 formed at least partially of a composite material; 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 (“FPF”), wherein

$\mathrm{FPF}=\left[\frac{c}{0.15·D}\right]/\left[\left[\frac{\sqrt{\mathrm{FPR}-1}}{\sqrt{0.4}}\right]/M_{\mathrm{tip},c}^{\left(\mathrm{RL}\right)}\right]-1.23;$ $\mathrm{wherein}$ $m_1·\left[M_{\mathrm{tip},c}^{\left(\mathrm{RL}\right)}-1.1\right]+9.14>\mathrm{FPF}>m_2·\left[M_{\mathrm{tip},c}^{\left(\mathrm{RL}\right)}-1.1\right],$

and wherein m 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 ML) 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 any preceding clause, wherein FPF is within a range equal to or greater than −0.8 and equal to or less than 8.4.

The turbomachine of any preceding clause, wherein FPF is within a range equal to or greater than 0 and equal to or less than 6.

The turbomachine of any preceding clause, wherein FPF is within a range equal to or greater than 1 and equal to or less than 2.

The turbomachine of any preceding clause, 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 any preceding clause, 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 any preceding clause, 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 any preceding clause, wherein FPR is within a range equal to or greater than 1.2 and equal to or less than 1.6.

The turbomachine of any preceding clause, wherein FPR is within a range equal to or greater than 1.3 and equal to or less than 1.5.

The turbomachine of any preceding clause, wherein FPR is within a range equal to or greater than 1.35 and equal to or less than 1.45.

The turbomachine of any preceding clause, 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 any preceding clause, 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 any preceding clause, wherein: the annular casing comprises an annular composite shell extending from a forward flange aft or downstream to an aft flange; and a circumferentially varying thickness portion of the annular casing includes the composite shell.

The turbomachine of any preceding clause, further comprising annular forward and aft seams between the annular composite shell and a composite back sheet.

The turbomachine of any preceding clause, further comprising the circumferentially varying thickness portion of the annular casing including aftwardly or downstream extending lobes or fingers of the composite back sheet on the composite shell in the aft seam, wherein a first thickness of the lobes or fingers is thicker than a second thickness of the composite shell circumferentially between the lobes or fingers.

The turbomachine of any preceding clause, further comprising an annular composite back sheet spaced radially outwardly of the annular composite shell and an annular filler layer disposed radially therebetween.

The turbomachine of any preceding clause, further comprising the filler layer including honeycomb.

The turbomachine of any preceding clause, further comprising annular forward and aft seams between the annular composite shell and composite back sheet axially trapping the filler layer between the forward and aft seams.

The turbomachine of any preceding clause, further comprising the circumferentially varying thickness portion of the annular casing including aftwardly or downstream extending lobes or fingers of the composite back sheet on the composite shell in the aft seam, wherein a first thickness of the lobes or fingers is thicker than a second thickness of the composite shell circumferentially between the lobes or fingers.

The turbomachine of any preceding clause, further comprising a metallic fan casing aft or downstream of and bolted to a composite fan case, the metallic fan casing including a metallic fan casing flange bolted to the aft flange.

A turbomachine comprising: an annular casing formed at least partially of a composite material; 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 fan tip Mach number (“M_tip,c^(RL)”) according to a Second Performance Factor (“SPF”), wherein

$\mathrm{SPF}=\sqrt{\frac{\pi }{4}\left(1-{\mathrm{HTR}}^2\right)/\left(\frac{\mathrm{BC}}{20}\right)}/\left(\frac{\sqrt{\mathrm{FPR}-1}}{\sqrt{0.4}}\right)/M_{\mathrm{tip},c}^{\left(\mathrm{RL}\right)}-0.97,$ $\mathrm{wherein}$ $m_3·\left[M_{\mathrm{tip},c}^{\left(\mathrm{RL}\right)}-1.1\right]+2.52>\mathrm{SPF}>m_4·\left[M_{\mathrm{tip},c}^{\left(\mathrm{RL}\right)}-1.1\right]$

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 any preceding clause, wherein SPF is within a range equal to or greater than 0.087 and equal to or less than 2.4.

The turbomachine of any preceding clause, wherein SPF is within a range equal to or greater than 1 and equal to or less than 2.

The turbomachine of any preceding clause, 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 any preceding clause, 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 any preceding clause, wherein HTR is within a range equal to or greater than 0.2 and equal to or less than 0.4.

The turbomachine of any preceding clause, wherein HTR is within a range equal to or greater than 0.25 and equal to or less than 0.35.

The turbomachine of any preceding clause, wherein FPR is within a range equal to or greater than 1.2 and equal to or less than 1.6.

The turbomachine of any preceding clause, wherein FPR is within a range equal to or greater than 1.3 and equal to or less than 1.5.

The turbomachine of any preceding clause, wherein FPR is within a range equal to or greater than 1.35 and equal to or less than 1.45.

The turbomachine of any preceding clause, wherein BC is within a range equal to or greater than 3 and equal to or less than 18.

The turbomachine of any preceding clause, wherein BC is within a range equal to or greater than 10 and equal to or less than 16.

The turbomachine of any preceding clause, 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 any preceding clause, 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 for an aircraft comprising: an annular casing formed at least partially of a composite material; 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 (“FPF”), wherein

$\mathrm{FPF}=\left[\frac{c}{0.15·D}\right]/\left[\left[\frac{\sqrt{\mathrm{FPR}-1}}{\sqrt{0.4}}\right]/M_{\mathrm{tip},c}^{\left(\mathrm{RL}\right)}\right]-1.23,$ $\mathrm{wherein}$ $m_1·\left[M_{\mathrm{tip},c}^{\left(\mathrm{RL}\right)}-1.1\right]+9.14>\mathrm{FPF}>m_2·\left[M_{\mathrm{tip},c}^{\left(\mathrm{RL}\right)}-1.1\right],$

and 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; wherein the fan includes a fan hub-to-tip ratio (“HTR”), a fan blade count (“BC”), the fan pressure ratio (“FPR”), and the redline corrected fan tip Mach number (“M_tip,c^(RL)”) according to a Second Performance Factor (“SPF”), wherein

$\mathrm{SPF}=\sqrt{\frac{\pi }{4}\left(1-{\mathrm{HTR}}^2\right)/\left(\frac{\mathrm{BC}}{20}\right)}/\left(\frac{\sqrt{\mathrm{FPR}-1}}{\sqrt{0.4}}\right)/M_{\mathrm{tip},c}^{\left(\mathrm{RL}\right)}-0.97,$ $\mathrm{wherein}$ $m_3·\left[M_{\mathrm{tip},c}^{\left(\mathrm{RL}\right)}-1.1\right]+2.52>\mathrm{SPF}>m_4·\left[M_{\mathrm{tip},c}^{\left(\mathrm{RL}\right)}-1.1\right]$

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 any preceding clause, wherein FPF is within a range equal to or greater than −0.8 and equal to or less than 8.4.

The turbomachine of any preceding clause, wherein FPF is within a range equal to or greater than 0 and equal to or less than 6.

The turbomachine of any preceding clause, wherein FPF is within a range equal to or greater than 1 and equal to or less than 2.

The turbomachine of any preceding clause, wherein SPF is within a range equal to or greater than 0.087 and equal to or less than 2.4.

The turbomachine of any preceding clause, wherein SPF is within a range equal to or greater than 1 and equal to or less than 2.

The turbomachine of any preceding clause, 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 any preceding clause, 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 any preceding clause, 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 any preceding clause, wherein HTR is within a range equal to or greater than 0.2 and equal to or less than 0.4.

The turbomachine of any preceding clause, wherein HTR is within a range equal to or greater than 0.25 and equal to or less than 0.35.

The turbomachine of any preceding clause, wherein FPR is within a range equal to or greater than 1.2 and equal to or less than 1.6.

The turbomachine of any preceding clause, wherein FPR is within a range equal to or greater than 1.3 and equal to or less than 1.5.

The turbomachine of any preceding clause, wherein FPR is within a range equal to or greater than 1.35 and equal to or less than 1.45.

The turbomachine of any preceding clause, wherein BC is within a range equal to or greater than 3 and equal to or less than 18.

The turbomachine of any preceding clause, wherein BC is within a range equal to or greater than 10 and equal to or less than 16.

The turbomachine of any preceding clause, 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 any preceding clause, 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.

Claims

1. A turbomachine for an aircraft comprising: an annular casing formed at least partially of a composite material; anda 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 (“Mtip,c(RL)”) according to a First Performance Factor (“FPF”),wherein

2. The turbomachine of claim 1, wherein: FPF is within a range equal to or greater than −0.8 and equal to or less than 8.4;Mtip,c(RL) is within a range equal to or greater than 0.8 and equal to or less than 1.5;ratio c/D is within a range equal to or greater than 0.1 and equal to or less than 0.3; andFPR is within a range equal to or greater than 1.2 and equal to or less than 1.6.

3. The turbomachine of claim 1, 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.

4. The turbomachine of claim 1, 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.

5. The turbomachine of claim 1, wherein: the annular casing comprises an annular composite shell extending from a forward flange aft or downstream to an aft flange; anda circumferentially varying thickness portion of the annular casing includes the composite shell.

6. The turbomachine of claim 5, further comprising annular forward and aft seams between the annular composite shell and a composite back sheet.

7. The turbomachine of claim 6, further comprising the circumferentially varying thickness portion of the annular casing including aftwardly or downstream extending lobes or fingers of the composite back sheet on the composite shell in the aft seam, wherein a first thickness of the lobes or fingers is thicker than a second thickness of the composite shell circumferentially between the lobes or fingers.

8. The turbomachine of claim 5, further comprising an annular composite back sheet spaced radially outwardly of the annular composite shell and an annular filler layer disposed radially therebetween.

9. The turbomachine of claim 8, further comprising the filler layer including honeycomb.

10. The turbomachine of claim 8, further comprising annular forward and aft seams between the annular composite shell and composite back sheet axially trapping the filler layer between the forward and aft seams.

11. The turbomachine of claim 10, further comprising the circumferentially varying thickness portion of the annular casing including aftwardly or downstream extending lobes or fingers of the composite back sheet on the composite shell in the aft seam, wherein a first thickness of the lobes or fingers is thicker than a second thickness of the composite shell circumferentially between the lobes or fingers.

12. The turbomachine of claim 5, further comprising a metallic fan casing aft or downstream of and bolted to a composite fan case, the metallic fan casing including a metallic fan casing flange bolted to the aft flange.

13. A turbomachine comprising: an annular casing formed at least partially of a composite material; anda 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 fan tip Mach number (“Mtip,c(RL)”) according to a Second Performance Factor (“SPF”),wherein

14. The turbomachine of claim 13, wherein: SPF is within a range equal to or greater than 0.087 and equal to or less than 2.4;Mtip,c(RL) is within a range equal to or greater than 0.8 and equal to or less than 1.5;HTR is within a range equal to or greater than 0.2 and equal to or less than 0.4;FPR is within a range equal to or greater than 1.2 and equal to or less than 1.6; andBC is within a range equal to or greater than 3 and equal to or less than 18.

15. The turbomachine of claim 13, 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.

16. The turbomachine of claim 13, 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.

17. A turbomachine for an aircraft comprising: an annular casing formed at least partially of a composite material; anda 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 (“Mtip,c(RL)”) according to a First Performance Factor (“FPF”), wherein

18. The turbomachine of claim 17, wherein: FPF is within a range equal to or greater than −0.8 and equal to or less than 8.4;SPF is within a range equal to or greater than 0.087 and equal to or less than 2.4;Mtip,c(RL) is within a range equal to or greater than 0.8 and equal to or less than 1.5;ratio c/D is within a range equal to or greater than 0.1 and equal to or less than 0.3;HTR is within a range equal to or greater than 0.2 and equal to or less than 0.4;FPR is within a range equal to or greater than 1.2 and equal to or less than 1.6; andBC is within a range equal to or greater than 3 and equal to or less than 18.

19. The turbomachine of claim 17, 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.

20. The turbomachine of claim 17, 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.