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 the second performance factor.

Description

[0001] This application is a divisional application of the invention patent application with application number 202311503952.0 filed on November 13, 2023 and invention name “Turbine and Assembly Method”. Technical Field

[0002] The present disclosure relates generally to jet engines and, more particularly, to jet engine fans. Background Art

[0003] 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.

[0004] 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

[0005] 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:

[0006] Figure 1 is a schematic cross-sectional view of a gas turbine engine according to an exemplary aspect of the present disclosure;

[0007] Figure 2 is a cross-sectional view of a fan blade according to an exemplary aspect of the present disclosure;

[0008] 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

[0009] 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.

[0010] 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

[0011] 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.

[0012] 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.

[0013] 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.

[0014] 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.

[0015] The term "radial" or "radially" refers to directions away from a common center. For example, in the overall 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.

[0016] The singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise.

[0017] 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.

[0018] 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.

[0019] 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 speed 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 speed of the turbine engine) that identify the number of fan blades required to produce improved results in terms of engine efficiency.

[0020] Various aspects of the present disclosure describe various 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 improved engine aerodynamic efficiency and / or improved 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. In one aspect, a first performance factor (FPF) of the fan module is given by the inequality In another aspect, the second performance factor (SPF) of the fan module is expressed by the inequality express.

[0021] 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.

[0022] 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 .

[0023] 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.

[0024] 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.

[0025] Still reference Figure 1In 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.

[0026] 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.

[0027] 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.

[0028] 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.

[0029] 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.

[0030] 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.

[0031] 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.

[0032] 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).

[0033] 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.

[0034] 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 speed. 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.

[0035] 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.

[0036] 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.

[0037] 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 fan tip speed "U" are based on the following relationship: c(tip) ", which is referred to herein as the first performance factor (FPF) of the fan module:

[0038]

[0039] 0.24*U c(tip) +489 > FPF >0.24*U c(tip) -12 +dy 1 (2)

[0040] Terminology 1The definition of is as follows: The ratio of the average fan chord "c" to the 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).

[0041] 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).

[0042] As used herein, “U c(tip) ” is the corrected fan tip speed at redline (e.g., the maximum allowable rotational speed of the fan at the redline shaft speed, which is directly coupled to the fan or through a reduction gearbox). The “fan tip speed” defined by the plurality of fan blades 40 refers to the linear speed of the outer tips of the fan blades 40 during operation of the fan 38. The corrected fan tip speed may be provided, for example, in ft / sec divided by an industry standard temperature correction. In an example approach, the corrected fan tip speed 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.

[0043] 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.

[0044] refer to Figure 3 The FPF value increases along the Y-axis, while the X-axis represents the red line corrected fan tip speed (U) increasing from left to right. c(tip) ). The FPF value falling on line 200 is equal to 0.24*U c(tip) –12. Figure 3 Eight example engine embodiments with unique FPF values ​​are shown. They are engines 210, 212, 214, 216, which may be referred to as low-speed engine designs (as shown in sub-plot area 218), and engines 220, 222, 224, 226, which may be referred to as high-speed engine designs (as shown in sub-plot area 228).

[0045] 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 for a given U below line 200 (within plot area 242) c(tip) Compared to an engine with a FPF value of 200 above the line (within the plot area 240), a given U c(tip) 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 U c(tip) This advantage may become more pronounced (e.g., for the next generation of ultra-high bypass ratio engines); for example, the improvement in engine performance based on redline tip speed may have an effect greater than 0.24*U c(tip) –4.5, greater than 0.24*U c(tip) +4 or greater than 0.24*U c(tip) +14 FPF values ​​(these other examples are generally represented by dashed line 202). In other words, these fan chord relationships in terms of FPF values ​​(line 202 represents the lower limit) can be more generally represented as FPF>0.24U c(tip) -12 + dy 1 , where dy 1 Equal to 7.5, 16, 25. 1 The value can vary between 0 and 500, 0 and 7.5, 7.5 and 16, 16 and 25, or a value greater than 25 and less than 500.

[0046] In another aspect, the inventors have 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 fan tip speed "U c(tip) ", which is referred to herein as the second performance factor (SPF) of the fan module:

[0047]

[0048] 0.15*U c(tip) + 654 > SPF > 0.15*U c(tip) +153+dy 2 (4)

[0049] Terminology 2The definitions are as follows. 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).

[0050] 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).

[0051] “FPR” and “U c(tip) ” refer to the fan pressure ratio and the corrected fan tip speed at the redline, respectively, as discussed above with respect to the average fan chord relationship. In this manner, FPR and / or U c(tip) The values ​​of can be the same or similar to those discussed with respect to the average fan chord relationship.

[0052] refer to Figure 4 , the SPF value increases along the Y-axis, while the X-axis represents the red line corrected fan tip speed (U) increasing from left to right c(tip) ). As R / LU c(tip) The SPF value falling on line 300 is equal to 0.15U c(tip) +153. Figure 4 Six example engine embodiments with unique SPF values ​​are shown, which are engines 310, 312, 314, 316, which may be referred to as low-speed engine designs (as shown in sub-plot area 318), and engines 320, 322, 324, 326, which may be referred to as high-speed engine designs (as shown in sub-plot area 328).

[0053] 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 for a given U below line 300 (within plot area 342) c(tip) Compared to an engine with a SPF value of 100 above line 300 (within plot region 340), a given U c(tip)The engine with an SPF value of 314, 318, 324 and 328 can allow for a reduced number of fan blades. In this way, engines 314, 318, 324 and 328 can provide advantages over engines 310, 312, 320 and 322, such as reduced cost and weight. In some cases, as the SPF value increases and U c(tip) This advantage may become more pronounced (e.g., for the next generation of ultra-high bypass ratio engines); for example, the improvement in engine performance based on redline tip speed may have an effect greater than 0.15*U c(tip) +158, greater than 0.15*U c(tip) +163 or greater than 0.15*U c(tip) +168 SPF value (these other examples are generally represented by dashed line 302). In other words, these fan blade number relationships in terms of SPF value (line 302 represents its lower limit) can be more generally represented as SPF>0.15U c(tip) +153+dy 2 , where dy 2 Equal to, for example, 5, 10, 15. 2 The value may vary between 0 and 500, 0 and 15, or may be a value greater than 15 and less than 500.

[0054] In this manner, using fan parameters such as fan pressure ratio, corrected fan tip speed, 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.

[0055] 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.

[0056] 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.

[0057] 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.

[0058] Further aspects are provided by the subject matter of the following clauses:

[0059] 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, a diameter ("D") of the fan, a fan pressure ratio ("FPR"), and a redline corrected fan tip speed ("U") according to a first performance factor. c(tip) ");in And 0.24*U c(tip) +489>FPF>0.24*U c(tip) –12+dy 1 , and where 0 <dy 1 <500.

[0060] 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.

[0061] 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.

[0062] A turbine according to one or more of these clauses, wherein the ratio c / D is equal to or greater than 0.1.

[0063] A turbine according to one or more of these clauses, wherein the ratio c / D is equal to or less than 0.3.

[0064] 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.

[0065] 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.

[0066] The turbine according to one or more of these clauses, wherein the FPR is equal to or greater than 1.2.

[0067] The turbine according to one or more of these clauses, wherein the FPR is equal to or greater than 1.3.

[0068] The turbine according to one or more of these clauses, wherein the FPR is equal to or less than 1.5.

[0069] The turbine of one or more of these clauses, wherein the FPR is equal to or less than 1.45.

[0070] 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.

[0071] 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.

[0072] A turbine according to one or more of these clauses, wherein dy 1 7.5 or more and 500 or less.

[0073] A turbine according to one or more of these clauses, wherein dy 1 Equal to or greater than 16 and equal to or less than 500.

[0074] A turbine according to one or more of these clauses, wherein dy 1 Equal to or greater than 25 and equal to or less than 500.

[0075] 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 fan tip speed ("U c(tip) ");in And 0.15*U c(tip) +654>SPF>0.15*U c(tip) +153+dy 2 , and where 0 <dy 2 <500.

[0076] 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.

[0077] 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.

[0078] 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.

[0079] The turbine according to one or more of these clauses, wherein the HTR is equal to or greater than 0.1.

[0080] A turbine according to one or more of these clauses, wherein the HTR is equal to or less than 0.5.

[0081] 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.

[0082] 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.

[0083] 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.

[0084] The turbine according to one or more of these clauses, wherein the FPR is equal to or greater than 1.2.

[0085] The turbine according to one or more of these clauses, wherein the FPR is equal to or greater than 1.3.

[0086] The turbine according to one or more of these clauses, wherein the FPR is equal to or less than 1.5.

[0087] The turbine of one or more of these clauses, wherein the FPR is equal to or less than 1.45.

[0088] 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.

[0089] 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.

[0090] A turbine according to one or more of these clauses, wherein dy 2 Equal to 5 and equal to or less than 500.

[0091] A turbine according to one or more of these clauses, wherein dy 2 Equal to 10 and equal to or less than 500.

[0092] A turbine according to one or more of these clauses, wherein dy 2 Equal to 15 and equal to or less than 500.

[0093] An assembly method, 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 speed ("U c(tip) "),in: And 0.24*U c(tip) +489>FPF>0.24*U c(tip) –12+dy 1 , and where 0 <dy 1 <500; or wherein the fan comprises a fan hub-to-tip ratio ("HTR"), a fan blade count ("BC"), a FPR, and a U according to a second performance factor ("SPF"). c(tip) ,in: And 0.15*U c(tip) +654>SPF>0.15*U c(tip) +153+dy 2 , and where 0 <dy 2 <500.

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 speed ("U c(tip) ”); in and Where 0.15*U c(tip) +654>SPF>0.15*U c(tip) +153+dy2, and 0 <dy2<500。 2. The turbine according to claim 1, characterized in that in, The HTR is in the range of equal to or greater than 0.1 and equal to or less than 0.

5.

3. The turbine according to claim 1, characterized in that in, BC is in the range of 10 or more and 18 or less.

4. The turbine according to claim 1, characterized in that in, The FPR is in the range of equal to or greater than 1.2 and equal to or less than 1.

5.

5. The turbine according to claim 1, characterized in that in, 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.

6. The turbine according to claim 1, characterized in that in, 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.

7. The turbine according to claim 1, characterized in that in, dy2 is equal to or greater than 5 and equal to or less than 500.

8. The turbine according to claim 1, characterized in that in, dy2 is equal to or greater than 10 and equal to or less than 500.

9. The turbine according to claim 1, characterized in that in, dy2 is equal to or greater than 15 and equal to or less than 500.

10. An assembly method, characterized in that: include: a fan mounted inside the annular housing for rotation 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 FPR, and a U c(tip) ,in: and 0.15*U c(tip) +654>SPF>0.15*U c(tip) +153+dy2, and 0 <dy2<500。

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