Turbomachine and method of assembly

Abstract

A turbomachine includes an annular housing and a fan disposed inside the annular housing and mounted for rotation about an axial centerline. The fan includes fan blades extending radially outward toward the annular housing. The fan has an average chord fan width according to a first performance factor. The fan has a number of fan blades according to a second performance factor.

Description

[0001] This application is a divisional application of the invention patent application with application number 202311503947.X filed on November 13, 2023 and invention name “Turbine and Assembly Method”.

[0002] CROSS-REFERENCE TO RELATED APPLICATIONS

[0003] This application is a continuation-in-part of application serial number 17 / 986,544 filed on November 14, 2022, which is incorporated herein by reference in its entirety. Technical Field

[0004] The present disclosure relates generally to jet engines and, more particularly, to jet engine fans. Background Art

[0005] In one form, a gas turbine engine may include a fan and a core arranged in flow communication with each other. The core generally includes a compressor section, a combustion section, a turbine section, and an exhaust section in a serial flow order. The fan and the core may be partially surrounded by an outer nacelle. In such a method, the outer nacelle and the core define a bypass airflow channel.

[0006] In operation, air is provided from the fan to the inlet of the compressor section, where one or more axial compressors gradually compress the air until it reaches the combustion section. Fuel is mixed with the compressed air and combusted using one or more fuel nozzles within the combustion section to provide combustion gases. The combustion gases are directed from the combustion section to the turbine section. The flow of combustion gases through the turbine section drives the turbine section and is then directed to be discharged to the atmosphere through the exhaust section. BRIEF DESCRIPTION OF THE DRAWINGS

[0007] A full and enabling disclosure of the present disclosure, including the best mode thereof, to one of ordinary skill in the art is set forth in the specification with reference to the accompanying drawings, in which:

[0008] Figure 1 is a schematic cross-sectional view of a gas turbine engine according to an exemplary aspect of the present disclosure;

[0009] Figure 2 is a cross-sectional view of a fan blade according to an exemplary aspect of the present disclosure;

[0010] Figure 3 showing a first example engine arranged on a first plot according to a first performance factor of a fan module according to the present disclosure; and

[0011] Figure 4 A second example engine is shown arranged on a second plot according to a second performance factor for a fan module according to the present disclosure.

[0012] The elements in the drawings are shown for simplicity and clarity and are not necessarily drawn to scale. For example, the size and / or relative position of some elements in the drawings may be exaggerated relative to other elements to help improve the understanding of the variations of the present disclosure. In addition, common but well-understood elements that are useful or necessary in commercially feasible embodiments are generally not described to facilitate less obstructed observation of these variations of the present disclosure. Certain actions and / or steps may be described or depicted in a specific order of occurrence, but those skilled in the art will understand that such specificity with respect to sequence is not actually required. DETAILED DESCRIPTION

[0013] 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.

[0014] 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. In addition, unless expressly stated otherwise, all embodiments described herein should be considered exemplary.

[0015] As used herein, the terms “first” and “second” may be used interchangeably to distinguish one component from another, and are not intended to indicate the position or importance of each component.

[0016] The terms "upstream" and "downstream" refer to relative directions with respect to flow in a path. For example, with respect to fluid flow, "upstream" refers to the direction from which the fluid is flowing, and "downstream" refers to the direction toward which the fluid is flowing.

[0017] The term "radial" or "radially" refers to directions away from a common center. For example, in the general context of turbine engines, radial refers to directions along a ray extending between a central longitudinal axis of the engine and the periphery of the engine.

[0018] As used herein, the term "composite material" refers to a material including non-metallic elements or materials. As used herein, "composite component" or "composite material" refers to a structure or component including any suitable composite material. Composite materials can be a combination of at least two or more non-metallic elements or materials. Examples of composite materials can be, but are not limited to, polymer matrix composites (PMC), ceramic matrix composites (CMC), metal matrix composites (MMC), carbon fibers, polymer resins, thermoplastic resins, bismaleimide (BMI) materials, polyimide materials, epoxy resins, glass fibers, and silicon matrix materials. Composite components (e.g., composite airfoils) can include several layers or several ply composite materials. The stiffness, material, and size of the layers or plies can be different to achieve a composite part of a desired composite component or component with a predetermined weight, size, stiffness, and strength. One or more layers of adhesive can be used to form or connect composite components. Adhesives can include resins and phenolic resins, wherein the adhesive may need to be cured at elevated temperatures or other hardening techniques.

[0019] As used herein, "PMC" refers to a class of materials. As an example, PMC materials are partially defined by prepregs, which are reinforcing materials pre-impregnated with polymer matrix materials (e.g., thermoplastic resins). Non-limiting examples of processes for producing thermoplastic prepregs include: hot melt prepregs, in which fiber reinforcements are pulled through a molten bath of resin; and powder prepregs, in which resins are deposited on fiber reinforcements, as a non-limiting example, electrostatically deposited on fiber reinforcements, and then adhered to fibers, as a non-limiting example, in an oven or with the help of heated rollers. Prepregs can be in the form of unidirectional tapes or woven fabrics, which are then stacked on top of each other to form the desired number of stacked plies of a part. Multilayer prepregs can be stacked to the appropriate thickness and orientation of a composite component, and then the resin is cured and solidified to provide a fiber-reinforced composite part. The resins used for PMC matrix materials can generally be classified as thermosetting resins or thermoplastic resins. Thermoplastic resins are generally classified as polymers that can soften and flow repeatedly when heated and can harden due to physical changes rather than chemical changes when fully cooled. Famous examples of thermoplastic resins include nylon, thermoplastic polyesters, polyaryletherketones and polycarbonate resins. Specific examples of high-performance thermoplastic resins that have been envisioned for aerospace applications include polyetheretherketone (PEEK), polyetherketoneketone (PEKK), polyetherimide (PEI), polyaryletherketone (PAEK) and polyphenylene sulfide (PPS). On the contrary, once fully cured into a hard rigid solid, thermosetting resins do not experience significant softening when heated, but thermally decompose when fully heated. Famous examples of thermosetting resins include epoxy resins, bismaleimide (BMI) and polyimide resins.

[0020] Instead of using prepreg, in another non-limiting example, by using thermoplastic polymers, braided fabrics can be utilized. Braided fabrics may include, but are not limited to, dry carbon fibers woven together with thermoplastic polymer fibers or filaments. Non-prepreg braided architectures can be made in a similar manner. In this way, the fiber volume of a part can be customized by specifying the relative concentrations of thermoplastic fibers and reinforcing fibers that have been woven or braided together. In addition, different types of reinforcing fibers can be braided or braided together at different concentrations to customize the properties of a part. For example, glass fibers, carbon fibers, and thermoplastic fibers can all be woven together at different concentrations to customize the properties of a part. Carbon fibers provide the strength of the system, glass fibers can be incorporated to enhance impact properties, which is a design feature of a part located near the engine inlet, and thermoplastic fibers provide bonding for reinforcing fibers.

[0021] In yet another non-limiting example, resin transfer molding (RTM) can be used to form at least a portion of a composite component. Typically, RTM includes applying a dry fiber or matrix material to a mold or cavity. The dry fiber or matrix material may include prepregs, woven materials, braided materials, or any combination thereof.

[0022] The resin may be pumped 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 is then cured and removed from the mold. When removed from the mold, the composite part may require a post-curing process.

[0023] It is contemplated that RTM may be a vacuum assisted process. That is, the air in the cavity or mold may be removed and replaced with resin prior to heating or curing. It is further contemplated that placement of dry fiber or matrix material may be manual or automated.

[0024] The dry fiber or matrix material may be shaped to form the composite part or to guide the resin. Optionally, additional layers or reinforcements of a material different from the dry fiber or matrix material may also be included or added prior to heating or curing.

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

[0026] The inventors sought to maximize the efficiency of a turbine engine during flight propulsion of an aircraft, and correspondingly reduce fuel consumption. In particular, the inventors focused on how the fan of a ducted turbine engine could be improved. Given several different proposed engine architectures, the inventors considered how the fan module would need to be changed to achieve mission requirements, and / or how the fan module could improve existing engine efficiency and / or fuel consumption. The inventors studied several engine architectures and then determined how the number of fan blades used by the fan, the average chord width of the fan blades, the diameter of the fan blades, the fan pressure ratio, the fan tip speed, and the hub-to-tip ratio affect engine efficiency and / or fuel consumption.

[0027] The inventors have discovered that in some engines, too many fan blades may add unnecessary cost to the engine design without providing a clear benefit, and may also add unnecessary weight to the aircraft, thereby reducing overall fuel efficiency (e.g., due to increased fuel burn). However, it has been discovered that a reduction in the number of fan blades may result in a reduction in the total fan blade area desired for effective propulsion, fan aeromechanical stability, and operability. The inventors have considered increasing the width of the fan blades, or the fan chord, to achieve the desired fan blade area with a lower number of fan blades. These considerations have been found to be of particular interest when the engine has a higher bypass ratio (i.e., a lower fan pressure ratio), and when the engine has a lower blade tip speed.

[0028] Determining the number of fan blades and the average fan chord to achieve a desired efficiency typically requires a time-consuming iterative process. As explained in more detail below, after evaluating many turbine engine architectures having different numbers of fan blades and average fan chords, it was unexpectedly discovered that there are certain relationships between the fan or fan blade diameter, the fan pressure ratio, and the redline-corrected fan tip Mach number of the turbine engine that identify the average fan chord required to produce improved results in terms of engine efficiency. It was further unexpectedly discovered that there are certain relationships between turbine engine parameters (including the hub-to-tip ratio of the fan, the fan pressure ratio, and the redline-corrected fan tip Mach number of the turbine engine) that identify the number of fan blades required to produce improved results in terms of engine efficiency.

[0029] 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 number of blades, which are believed to result in increased engine aerodynamic efficiency and / or increased fuel efficiency. According to the present disclosure, a turbine for powering an aircraft in flight includes an annular casing and a fan disposed within the annular casing and mounted for rotation about an axial centerline, the fan including fan blades extending radially outwardly toward the annular casing.

[0030] In one aspect, the first performance factor (FPF) of the fan module is expressed according to the following inequality:

[0031]

[0032] On the other hand, the second performance factor (SPF) of the fan module is expressed by the following inequality:

[0033]

[0034] Reference will now be made in detail to embodiments of the present 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 present disclosure, and together with the description, serve to explain the principles of the present disclosure. The detailed description uses numerical and letter designations to refer to features in the drawings. Like or similar names in the drawings and description have been used to refer to like or similar parts of the present disclosure.

[0035] Referring now to the drawings, wherein like numerals refer to like elements throughout, Figure 1 is a schematic cross-sectional view of a turbomachine, and more particularly a gas turbine engine, according to an exemplary embodiment of the present disclosure. More specifically, Figure 1 In the embodiment of the present invention, the gas turbine engine is a high bypass turbofan jet engine 10, referred to herein as a "turbofan jet engine 10". Figure 1 As shown, the turbofan jet engine 10 defines an axial direction A (extending parallel to a longitudinal centerline 12 provided for reference) and a radial direction R. Generally, the turbofan jet engine 10 includes a fan section 14 and a core turbine engine 16 disposed downstream of the fan section 14 .

[0036] The depicted exemplary core turbine engine 16 generally includes a substantially tubular casing 18 defining an annular inlet 20. The casing 18 encloses, in serial flow relationship: a compressor section including a supercharger 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 an ejection exhaust nozzle section 32. A high pressure (HP) shaft or spool 34 drivingly connects the high pressure turbine 28 to the high pressure compressor 24. A low pressure (LP) shaft or spool 36 drivingly connects the low pressure turbine 30 to the low pressure compressor 22.

[0037] The fan blades 40 extend outwardly from the disk 42 generally along a radial direction R. For the depicted embodiment, 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 manner. One or more of the fan blades 40 can be rotated relative to the disk 42 about a pitch axis P by means of the fan blades 40 being operably coupled to a suitable actuating member 44, the actuating member 44 being configured to generally change the pitch of the fan blades 40 in unison together. In some methods, the fan is a fixed pitch fan, and there is no actuating member 44. The fan blades 40, the disk 42, and the actuating member 44 can be rotatable about the longitudinal centerline 12 together through the low pressure spool 36 across the power gearbox 46. The power gearbox 46 includes a plurality of gears for reducing the rotational speed of the low pressure spool 36 to a more efficient rotating fan speed. In some methods, the low pressure spool 36 can directly drive the fan without the power gearbox 46.

[0038] The power gearbox 46 may include a plurality of gears, including an input and an output, and may also include one or more intermediate gears disposed between the input and the output and / or interconnecting the input and the output. The input may include a first speed, and the output may have a second speed. In some examples, the gear ratio of the first speed to the second speed is equal to or greater than 3.2 and equal to or less than 5.0. The power gearbox 46 may include various types and / or configurations. In some examples, the power gearbox 46 is a single-stage gearbox. In other examples, the power gearbox 46 is a multi-stage gearbox. In some examples, the power gearbox 46 is an epicyclic gearbox. In some examples, the power gearbox 46 is a non-episodic gearbox (e.g., a compound gearbox). More specifically, in some cases, the power gearbox 46 is an epicyclic gearbox configured as a planetary gear configuration. The planetary gear configuration includes a sun gear, a plurality of planetary gears (also referred to as "planetary gears"), and a ring gear. The sun gear is an input and is coupled to a power turbine (e.g., a low-pressure turbine) such that the sun gear and the power turbine rotate at the same speed. The star gear is disposed between the sun gear and the ring gear and interconnects the sun gear and the ring gear. The star gear is rotatably coupled to a fixed bracket. Therefore, the star gears can rotate about their respective axes, but cannot orbit together relative to the sun gear or the ring gear. As another example, the power gearbox 46 is an epicyclic gearbox configured as a planetary gear structure. The planetary gear structure includes a sun gear, a plurality of planetary gears and a ring gear. The sun gear is an input and is coupled to a power turbine. The planetary gears are disposed between the sun gear and the ring gear and interconnect the sun gear and the ring gear. The planetary gears are rotatably coupled to a rotatable bracket. In this way, the planetary gears can rotate about their respective axes and also rotate together with the bracket relative to the sun gear and the ring gear. The bracket is an output and is coupled to a fan assembly. The ring gear is fixed and cannot rotate.

[0039] Still reference Figure 1 In an exemplary embodiment of the present invention, the disk 42 is covered by a rotatable forward hub 48 that is aerodynamically shaped to facilitate airflow through the plurality of fan blades 40. In addition, the exemplary fan section 14 includes an annular fan case or outer nacelle 50 that circumferentially surrounds at least a portion of the variable pitch fan 38 and / or the core turbine engine 16. It should be understood that the outer nacelle 50 can be configured to be supported relative to the core turbine engine 16 by a plurality of circumferentially spaced outlet guide vanes 52. In addition, a downstream section 54 of the outer nacelle 50 can extend over an outer portion of the core turbine engine 16 so as to define a bypass airflow passage 56 therebetween.

[0040] During operation of the turbofan jet engine 10, a quantity of air 58 enters the turbofan jet engine 10 through the outer nacelle 50 and / or the associated inlet 60 of the fan section 14. As the quantity of air 58 passes through the fan blades 40, a first portion 62 of the air 58, as indicated by arrow 62, is directed or directed into the bypass airflow passage 56, and a second portion 64 of the air 58, as indicated by arrow 64, is directed or directed into the low-pressure compressor 22. The ratio between the first portion 62 of the air 58 and the second portion 64 of the air 58 is generally referred to as the bypass ratio. The pressure of the second portion 64 of the air 58 is then increased when the second portion 64 of the air 58 is directed through the high-pressure (HP) compressor 24 and enters the combustion section 26, where the second portion 64 of the air 58 is mixed with fuel and combusted to provide combustion gases 66. The combustion gases 66 are then directed through the high-pressure turbine 28 and the low-pressure turbine 30, where a portion of the heat energy and / or kinetic energy is extracted from the combustion gases 66.

[0041] The combustion gases 66 are then directed through the jet exhaust nozzle section 32 of the core turbine engine 16 to provide propulsive thrust. At the same time, the pressure of the first portion 62 of the air 58 is significantly increased as it is directed through the bypass airflow passage 56 before being discharged from the fan nozzle exhaust section 76 of the turbofan jet engine 10, also providing propulsive thrust.

[0042] However, it should be understood that Figure 1The turbofan jet engine 10 depicted in the embodiment is merely an example, and in other exemplary embodiments, aspects of the present disclosure may be additionally or alternatively applied to any other suitable gas turbine engine. For example, in other exemplary embodiments, the turbofan jet engine 10 may alternatively be any other suitable aviation gas turbine engine, such as a turbojet engine, a turboshaft engine, a turboprop engine, etc. In addition, in still other exemplary embodiments, the turbofan jet engine 10 may include or be operably connected to any other suitable accessory system. Additionally or alternatively, the exemplary turbofan jet engine 10 may not include or be operably connected to one or more of the accessory systems discussed above.

[0043] The fan blades 40 of the jet engine 10 can be made of PMC material with a metal leading edge to protect the airfoil from foreign objects (such as bird strikes). When the airfoil is affected by flutter during operation, the polymer matrix composite (PMC) material used for the airfoil can be more durable and / or show improved performance. In some embodiments, an engine with fewer fan blades (e.g., less than 25 fan blades) and a wider chord (c) (such as an engine with a blade count (BC) from 14 to 18 or from 16 to 20 fan blades and a chord to diameter ratio (c / D) greater than 0.17, or greater than 0.19, and less than 0.3 (e.g., less than 0.21)) has a fan blade airfoil made of PMC material with a metal leading edge.

[0044] Figure 2 is a cross-sectional view of fan blade 40 viewed radially (eg, toward the axis of rotation). First axis 100 is parallel to Figure 1 The second axis 102 is parallel to the circumferential direction θ.

[0045] The fan blade 40 includes a low pressure surface 110 and an opposing high pressure surface 112, which are respectively located at the proximal end 40a and the distal end 40b of the fan blade 40 (eg, Figure 1 The fan blade 40 also includes a leading edge 114 and a trailing edge 116.

[0046] Low pressure surface 110, high pressure surface 112, leading edge 114, and trailing edge 116 form profile 118 of fan blade 40. Profile 118 defines an average camber 120 extending from leading edge 114 to trailing edge 116 and equidistant from low pressure surface 110 and high pressure surface 112.

[0047] The profile 118 further defines a local chord 122 that represents the linear distance from the leading edge 114 to the trailing edge 116 (relative to a particular cross-section through the blade).

[0048] In some approaches, the fan blade 40 may have a profile 118 that varies along the radial height of the fan blade 40 between the proximal end 40a and the distal end 40b. For example, in some fan blade designs, the 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. Thus, the length of the chord 122 may vary along the radial height of the fan blade 40. In this manner, an average chord length may be derived for the fan blade that accounts for the variation in the chord length 122 along the radial height of the fan blade 40.

[0049] As previously mentioned, during the process of improving the fan module portion of various engine architectures, the inventors discovered the relationship between time scales (including fan pressure ratio, fan diameter, and corrected fan tip Mach number. More specifically, and as discussed in more detail below, the inventors have discovered the relationship between the ratio of the axial flow time scale to the rotational time scale and suitable parameters for implementing these relationships for the engine.

[0050] The aircraft turbine engine architecture developed by the inventor includes a fan module and an engine core as main components. The core includes one or more compressor stages and a turbine stage. The compressor stage generally includes a high-pressure compressor stage and a low-pressure compressor stage, and the turbine similarly includes a high-pressure stage and a low-pressure stage. The fan module that provides improved efficiency is not independent of these other parts of the engine, because when a part is improved or modified, there is always a trade-off. The improved efficiency brought by the fan can be reflected in weight reduction, installation resistance reduction, load balancing or management (dynamic or static load), aerodynamic efficiency / fan interaction with output guide vanes through the fan duct, and other factors. In order to improve what the fan can provide (the positive benefits of fan design), other parts of the engine often need to make sacrifices (the negative benefits of fan design). Or, when viewed independently of a specific core design or frame requirements, considering the impact that this fan design will have on other parts of the engine (such as compressor operating margin, fan and output guide vanes (OGV) and the balance of the power gearbox, and the position of the low-pressure compressor (packaging impact)), the benefits of the new fan design usually need to be modified or impractical. 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 (eg, distributed electric propulsion systems in which a gas turbine drives multiple fans).

[0051] In an effort to design an improved fan module, the inventors considered trade-offs between fan module improvements and other potential negative effects or limitations of fan module design, and unexpectedly discovered certain relationships that define the improved fan design, which are now described in detail.

[0052] In one aspect, the inventors have found that the average fan chord "c", the fan diameter "D" (e.g., the tip-to-tip dimension of the fan), the fan pressure ratio "FPR" of the fan, and the corrected redline fan tip Mach number are related by The following relationship is referred to herein as the first performance factor ("FPF") of the fan module:

[0053]

[0054] The ratio of average fan chord "c" to fan diameter "D" is a dimensionless chord-width ratio greater than 0.1 (eg, greater than 0.15, greater than 0.17, or greater than 0.19) and less than 0.3 (eg, less than 0.21).

[0055] As used herein, “fan pressure ratio” (FPR) refers to the ratio of the stagnation pressure immediately downstream of the plurality of exit guide vanes 52 during operation of the fan 38 to the stagnation pressure immediately upstream of the plurality of fan blades 40 during operation of the fan 38. The fraction may be used as a proxy for the ratio of increase in axial flow velocity through the fan. The fan pressure ratio is greater than 1.2 (eg, greater than 1.3) and less than 1.5 (eg, less than 1.45, less than 1.42, or less than 1.4).

[0056] As used herein, is the corrected fan tip Mach number at redline (e.g., the maximum allowable rotational speed of the fan at the redline shaft speed, which is coupled directly to the fan or through a reduction gearbox). “Fan tip speed” refers to the linear speed of the outer tips of fan blades 40 during operation of fan 38. “Corrected fan tip speed” (referred to as “U tip,c ”) can be provided, for example, in ft / sec divided by an industry standard temperature correction. In an example method, U tip,c It can 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. The “corrected fan tip Mach number” is the value obtained by changing U tip,c The dimensionless value is obtained by dividing by the accepted speed of sound under standard day sea level atmospheric conditions (i.e., 1,116.45 ft / sec). Thus, It may be less than 1.34 (eg, less than 1.12 or less than 0.99) and greater than 0.45.

[0057] As defined in (1), FPF can be considered to represent a ratio of velocities. When taking into account the normalized chord width "c / D", FPF can be considered to be the correlation of the time scale of blade rotation to the time it takes a flow particle to traverse the average chord length of the fan when the engine is operating under static conditions.

[0058] Refer to the inequality defined in (2) and Figure 3 , an example engine embodiment is shown with a characteristic FPF value and a corresponding redline corrected fan tip Mach number The values ​​of FPF increase along the Y-axis, while the X-axis represents the redline-corrected fan tip Mach number increasing from left to right. Figure 3 Also shown is a first line 200 and a second line 202 offset from the first line 200 along the Y axis. The first line 200 and the second line 202 are given by inequality (2): As used herein, "m1" refers to the slope of lines 200, 202, "1.1" refers to the reference corrected red line tip Mach number when the Y intercept is defined in the FPF, and Δy1 refers to the offset from the Y intercept along the Y axis.

[0059] like Figure 3 As shown, the first line 200 and the second line 202 are piecewise linear dividing curves; that is, depending on the The first line 200 and the second line 202 have different slopes "m1". More specifically, when When the value of is equal to or greater than 1.1, the first line 200 and the second line 202 have a slope "m1" equal to 0.87. When the value of is less than 1.1, the first line 200 and the second line 202 have a slope "m1" equal to 3.34. Although depicted as a piecewise linear division curve, the slow scaling is actually nonlinear and has a slope toward Figure 3 The lower portion of the plot reduces the advantage of the c / D design, which is associated with lower FPR and lower Associated.

[0060] As discussed, Δy1 refers to the 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 may be a value greater than 0.2 and less than 6.

[0061] Figure 3Eight example engine embodiments are shown, where engines 210, 212, 214, 216 may be referred to as low speed engine designs (as shown in sub-plot area 218), and engines 220, 222, 224, 226 may be referred to as high speed engine designs (as shown in sub-plot area 228). Each of engines 210, 212, 214, 216, 220, 222, 224, 226 has a gear ratio within the range of equal to or greater than 3.2 and less than or equal to 5.0.

[0062] As indicated by the FPF indicating a particular fan chord relationship, the inventors have discovered a limited or narrow selection of average fan chords that uniquely account for other factors associated with the fan and engine type. For example, the inventors have determined that with a given fan chord below line 200 (within plot area 242) Compared with the engine with a FPF value of 200 above the line (within the plot area 240), In this manner, engines 214, 216, 224, and 226 may provide advantages over engines 210, 212, 220, and 222, such as reduced fan blade count (discussed in more detail below), increased aeromechanical stability, and reduced fan lift coefficient C during takeoff of the aircraft. L In some cases, as FPF increases and This advantage may become more pronounced as the value of the bypass ratio decreases (e.g., for the next generation of ultra-high bypass ratio engines). For example, the improvement in engine performance based on the redline tip Mach number may have a value greater than Greater than Greater than Greater than or greater than 1.1]+0.2 FPF values ​​(these other examples are schematically represented by dashed line 202).

[0063] In another aspect, the inventors found that the fan blade count "BC", the fan's hub-to-tip ratio "HTR", the fan's fan pressure ratio "FPR", and the corrected redline fan tip Mach number are based on the following relationships: The following relationship is referred to herein as the second performance factor ("SPF") of the fan module:

[0064]

[0065] With respect to the hub to tip ratio "HTR", the fan blade defines a hub radius (R hub ) and tip radius (R tip), the hub radius is the radius of the leading edge at the hub relative to the centerline of the fan, and the tip radius is the radius of the leading edge at the 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).

[0066] The blade count "BC" corresponds to the number of fan blades arranged circumferentially around the fan hub. The blade count is between 10 fan blades and 40 fan blades. In certain example methods, the blade count is less than or equal to 18 fan blades (e.g., 16 or fewer fan blades).

[0067] "FPR" and are the fan pressure ratio and the redline-corrected fan tip Mach number, respectively, as discussed above with respect to the average fan chord relationship. In this way, FPR and The values ​​of one or more of may be the same as those discussed with respect to the average fan-chord relationship.

[0068] Refer to the inequality defined in (4) and Figure 4 , an example engine embodiment is shown with a characteristic SPF value and a corresponding redline corrected fan tip Mach number The SPF values ​​increase along the Y-axis, while the X-axis represents the redline corrected fan tip Mach number increasing from left to right. Figure 4 Also shown is a first line 300 and a second line 302 offset from the first line 300 along the Y axis. The first line 300 and the second line 302 are determined by inequality (4): Partially limited.

[0069] like Figure 4 As shown, the first line 300 and the second line 302 are piecewise linear dividing curves; that is, depending on the The first line 300 and the second line 302 have different slopes "m2". More specifically, when When the value of is equal to or greater than 1.1, the first line 300 and the second line 302 have a slope "m2" equal to 0.41. When the value of is less than 1.1, the first line 300 and the second line 302 have a slope “m2” equal to 0.55.

[0070] As used herein, "m2" refers to the slope of lines 300, 302, which is equal to 1 as shown. "1.1" refers to the reference corrected red line tip Mach number when the Y intercept is defined, and Δy2 refers to the 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 may be a value greater than 0.6 and less than 1.5.

[0071] Figure 4 Eight example engine embodiments are shown, where engines 310, 312, 314, 316 may be referred to as low speed engine designs (as shown in sub-plot area 318), and engines 320, 322, 324, 326 may be referred to as high speed engine designs (as shown in sub-plot area 328). Each of engines 310, 312, 314, 316, 320, 322, 324, 326 has a gear ratio within the range of equal to or greater than 3.2 and less than or equal to 5.0.

[0072] As indicated by the SPF indicating a particular fan blade count relationship, the inventors have discovered a limited or narrow selection of fan blade counts that uniquely accounts for other factors associated with the fan and engine type. For example, the inventors have determined that with a given fan blade count below line 300 (within plot area 342) Compared with the engine with an SPF value of 300 (within the plot area 340), An engine with an SPF value of 0.040 may allow for a reduced number of fan blades. In this manner, engines 314, 316, 324, and 326 may provide advantages over engines 310, 312, 320, and 322, such as reduced cost and weight. In some cases, as the SPF value increases and This advantage may become more pronounced as the value of the bypass ratio decreases (e.g., for the next generation of ultra-high bypass ratio engines). For example, the improvement in engine performance based on the redline tip Mach number may have a value greater than Greater than Greater than Greater than Greater than Greater than or greater than (these other examples are schematically represented by dashed line 302).

[0073] In this manner, using fan parameters such as fan pressure ratio, corrected fan tip Mach number, fan diameter, and hub-to-tip ratio, the inventors have discovered methods to obtain the average chord width using the above average fan chord relationship and the fan blade count using the above fan blade count relationship. These derived constraints guide one to select the fan chord width, blade count, or both that are appropriate for a particular engine architecture and mission requirements based on the engine's unique environment and tradeoffs in design (as described above), which is believed to result in an improved engine.

[0074] In another aspect, the FPF and SPF can also be used as a design tool to down-select or provide guidance to reduce the number of candidate designs for fan blade count and average fan chord to design a fan module for a particular architecture. In this way, it is known early in the design process what constraints or limitations the fan module will impose given the mission objectives, thereby improving the engine architecture overall.

[0075] In another aspect, an assembly method is provided. The method includes mounting a fan inside an annular housing for rotation about an axial centerline. The fan includes fan blades extending radially outward toward the annular housing. The fan also includes an average fan chord width based on the first performance factor ("FPF") and / or a number of fan blades based on the second performance factor ("SPF") described above.

[0076] This written description uses examples to disclose the present disclosure, including the best mode, and also enables any person skilled in the art to practice the present disclosure, including making and using any device or system and performing any combined methods. The patent scope of the present disclosure is defined by the claims, and may include other examples that occur to those skilled in the art. If these other examples include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements that do not differ substantially from the literal language of the claims, these other examples are intended to fall within the scope of the claims.

[0077] Further aspects are provided by the subject matter of the following clauses:

[0078] A turbomachine for an aircraft, comprising: an annular casing; and a fan, the fan being disposed inside the annular casing and mounted to rotate about an axial centerline, the fan comprising fan blades extending radially outwardly toward the annular casing; wherein the fan comprises an average fan chord width ("c") of the fan blades according to a first performance factor, a diameter ("D") of the fan, a fan pressure ratio ("FPR"), and a redline corrected fan tip Mach number in And among them And where 0<Δy1<6.

[0079] A turbine according to one or more of these clauses, wherein: In the range of equal to or greater than 0.45 and equal to or less than 1.34.

[0080] A turbine according to one or more of these clauses, wherein: In the range of equal to or greater than 0.45 and equal to or less than 1.12.

[0081] A turbine according to one or more of these clauses, wherein when When greater than or equal to 1.1, m1 is equal to 0.87.

[0082] A turbine according to one or more of these clauses, wherein when When greater than or equal to 1.1, m1 is equal to 1.0.

[0083] A turbine according to one or more of these clauses, wherein when When it is less than 1.1, m1 is equal to 2.5.

[0084] A turbine according to one or more of these clauses, wherein when When it is less than 1.1, m1 is equal to 3.34.

[0085] The turbine according to one or more of these clauses, wherein Δy1 is equal to or greater than 0.0125 and less than 6.

[0086] A turbine according to one or more of these clauses, wherein Δy1 is equal to or greater than 0.04 and less than 6.

[0087] The turbine according to one or more of these clauses, wherein Δy1 is equal to or greater than 0.07 and less than 6.

[0088] A turbine according to one or more of these clauses, wherein Δy1 is equal to or greater than 0.1 and less than 6.

[0089] A turbine according to one or more of these clauses, wherein Δy1 is equal to or greater than 0.2 and less than 6.

[0090] The turbine according to one or more of these clauses, wherein the turbine has a gear ratio in the range of equal to or greater than 3.2 and equal to or less than 5.0.

[0091] A turbine according to one or more of these clauses, wherein the fan blades comprise an airfoil portion made of a polymer matrix composite (PMC) material, a metal leading edge, the ratio c / D is between 0.16 and 0.21, and the fan has between 16 and 25 fan blades.

[0092] The turbine according to one or more of these clauses, wherein the ratio c / D is within the range of equal to or greater than 0.1 and equal to or less than 0.3.

[0093] The turbine according to one or more of these clauses, wherein the ratio c / D is in the range of equal to or greater than 0.15 and equal to or less than 0.21.

[0094] A turbine according to one or more of these clauses, wherein the ratio c / D is equal to or greater than 0.1.

[0095] A turbine according to one or more of these clauses, wherein the ratio c / D is equal to or less than 0.3.

[0096] The turbine according to one or more of these clauses, wherein the FPR is within the range of equal to or greater than 1.2 and equal to or less than 1.5.

[0097] The turbine according to one or more of these clauses, wherein the FPR is within the range of equal to or greater than 1.3 and equal to or less than 1.45.

[0098] The turbine according to one or more of these clauses, wherein the FPR is equal to or greater than 1.2.

[0099] The turbine according to one or more of these clauses, wherein the FPR is equal to or greater than 1.3.

[0100] The turbine according to one or more of these clauses, wherein the FPR is equal to or less than 1.5.

[0101] The turbine of one or more of these clauses, wherein the FPR is equal to or less than 1.45.

[0102] A turbine according to one or more of these clauses, wherein U c(tip) In the range of equal to or greater than 500 ft / sec and equal to or less than 1,500 ft / sec.

[0103] A turbine according to one or more of these clauses, wherein U c(tip) In the range of equal to or greater than 500 ft / sec and equal to or less than 1,250 ft / sec.

[0104] A turbomachine, comprising: an annular casing; and a fan, the fan being disposed inside the annular casing and mounted to rotate about an axial centerline, the fan comprising fan blades extending radially outwardly toward the annular casing; wherein the fan comprises 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 according to a second performance factor ("SPF") in And where 0<Δy2<1.5.

[0105] The turbine according to one or more of these clauses, wherein Δy2 is equal to or greater than 0.0075 and less than 1.5.

[0106] The turbine according to one or more of these clauses, wherein Δy2 is equal to or greater than 0.01 and less than 1.5.

[0107] The turbine according to one or more of these clauses, wherein Δy2 is equal to or greater than 0.02 and less than 1.5.

[0108] The turbine according to one or more of these clauses, wherein Δy2 is equal to or greater than 0.024 and less than 1.5.

[0109] The turbine according to one or more of these clauses, wherein Δy2 is equal to or greater than 0.037 and less than 1.5.

[0110] A turbine according to one or more of these clauses, wherein Δy2 is equal to or greater than 0.04 and less than 1.5.

[0111] A turbine according to one or more of these clauses, wherein Δy2 is equal to or greater than 0.06 and less than 1.5.

[0112] A turbine according to one or more of these clauses, wherein when When greater than or equal to 1.1, m2 is equal to 0.41.

[0113] A turbine according to one or more of these clauses, wherein when When it is less than 1.1, m2 is equal to 0.55.

[0114] The turbine according to one or more of these clauses, wherein the HTR is within the range of equal to or greater than 0.1 and equal to or less than 0.5.

[0115] The turbine according to one or more of these clauses, wherein the HTR is within the range of equal to or greater than 0.2 and equal to or less than 0.275.

[0116] The turbine according to one or more of these clauses, wherein the HTR is in the range of equal to or greater than 0.2 and equal to or less than 0.25.

[0117] The turbine according to one or more of these clauses, wherein the HTR is equal to or greater than 0.1.

[0118] A turbine according to one or more of these clauses, wherein the HTR is equal to or less than 0.5.

[0119] The turbine according to one or more of these clauses, wherein BC is in the range of equal to or greater than 10 and equal to or less than 18.

[0120] The turbine according to one or more of these clauses, wherein the FPR is within the range of equal to or greater than 1.2 and equal to or less than 1.5.

[0121] The turbine according to one or more of these clauses, wherein the FPR is within the range of equal to or greater than 1.3 and equal to or less than 1.45.

[0122] A method of assembling, comprising: mounting a fan inside an annular housing for rotation about an axial centerline, the fan comprising fan blades extending radially outwardly toward the annular housing; wherein the fan comprises an average fan chord width ("c") of the fan blades according to a first performance factor ("FPF"), a diameter ("D") of the fan, a fan pressure ratio ("FPR"), and a redline corrected fan tip redline Mach number in:

[0123] and 0<Δy1<6; or wherein the fan comprises a fan hub-to-tip ratio ("HTR"), a fan blade count ("BC"), a FPR and in:

[0124]

[0125] And 0

Claims

1. A turbine, characterized in that: include: annular housing; as well as a fan disposed within the annular housing and mounted to rotate about an axial centerline, the fan including fan blades extending radially outwardly toward the annular housing; 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 based on a second performance factor ("SPF"). in in and Where 0<Δy2<1.

5.

2. The turbine according to claim 1, characterized in that in, In the range of equal to or greater than 0.45 and equal to or less than 1.

34.

3. The turbine according to claim 1, characterized in that in, when When greater than or equal to 1.1, m2 is equal to 0.

41.

4. The turbine according to claim 1, characterized in that in, when When it is less than 1.1, m2 is equal to 0.

55.

5. The turbine according to claim 1, characterized in that in, Δy2 is equal to or greater than 0.0075 and less than 1.

5.

6. The turbine according to claim 1, characterized in that in, Δy2 is equal to or greater than 0.02 and less than 1.

5.

7. The turbine according to claim 1, characterized in that in, Δy2 is equal to or greater than 0.04 and less than 1.

5.

8. The turbine according to claim 1, characterized in that in, Δy2 is equal to or greater than 0.06 and less than 1.

5.

9. The turbine according to claim 1, characterized in that in, The turbine has a gear ratio within a range of equal to or greater than 3.2 and equal to or less than 5.

0.

10. The turbine according to claim 1, characterized in that in, The fan blade comprises an airfoil portion made of a polymer matrix composite (PMC) material, a metal leading edge, a ratio c / D between 0.16 and 0.21, and the fan has between 16 and 25 fan blades.

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