Variable pitch fans for turbomachinery engines

Abstract

A turbomachine engine includes a fan assembly having a plurality of variable pitch fan blades. The fan blades are configured such that they define a first VPF parameter and a second VPF parameter. The first VPF parameter is defined as a hub-to-tip radius ratio divided by a fan pressure ratio. The second VPF parameter is defined as a bearing spanwise force divided by a fan area. In some examples, the first VPF parameter is in a range of 0.1 to 0.25 and the second VPF parameter is in a range of 2-30 lbf / in 2 In other examples, the first VPF parameter is in the range of 0.1 to 0.4, and the second VPF parameter is in the range of 5.25-30 lbf / in 2 In some examples, the turbomachinery engine further includes a pitch changing mechanism, a blade assembly, a core engine, and a gearbox.

Description

Technical Field

[0001] The present disclosure relates generally to turbomachinery engines including fan assemblies, and more particularly to apparatus and methods directed to variable pitch fans for certain turbomachinery engine configurations. Background Art

[0002] A turbofan engine consists of a core engine that drives a bypass fan. The bypass fan generates most of the turbofan engine's thrust. The generated thrust can be used to move a payload (e.g., an aircraft).

[0003] In some examples, the turbofan engine is configured as a direct drive engine. A direct drive engine is configured such that the core engine's power turbine (e.g., a low-pressure turbine) is directly coupled to a bypass fan. Thus, the power turbine and the bypass fan rotate at the same rotational speed (i.e., the same rpm).

[0004] In other examples, a turbofan engine can be configured as a geared engine. A geared engine includes a gearbox disposed between and interconnecting the core engine's bypass fan and power turbine. For example, the gearbox allows the core engine's power turbine to rotate at a different speed than the bypass fan. Thus, for example, the gearbox can allow the core engine's power turbine and bypass fan to operate at their respective rotational speeds to achieve maximum efficiency and / or power generation.

[0005] In some examples, the propeller of a turbomachinery engine may be a fan enclosed within a fan case and / or nacelle. This type of turbomachinery engine may be referred to as a "bypass engine."

[0006] In other examples, the propeller of the turbomachinery engine may be exposed (eg, not within a fan case or nacelle). This type of turbomachinery engine may be referred to as an "open rotor engine" or "unducted engine."

[0007] In some examples, a turbofan engine may include a fixed-pitch fan, which is typical in commercial engines. In this configuration, the fan's pitch is static and configured to accommodate various engine operating conditions (e.g., takeoff, climb, cruise, approach, etc.).

[0008] In other examples, a turbofan engine may include a variable pitch fan. In such a configuration, the fan's pitch (or blade angle) can be adjusted as engine operating conditions change to improve propulsion efficiency.

[0009] Despite having certain advantages, engines that include a gearbox and / or a variable pitch fan may have one or more disadvantages. For example, including a gearbox and / or a variable pitch fan in a turbofan engine can introduce additional complexity to the engine. For example, this can make engine development and / or manufacturing more difficult. Therefore, there is a need for improved turbofan engines that include a gearbox and / or a variable pitch fan. There is also a need for apparatus and methods that can be used to more efficiently and / or more accurately develop and manufacture geared turbofan engines with variable pitch fans. Summary of the Invention

[0010] Aspects and advantages of the disclosed technology will be set forth in part in the following description, or may be obvious from the description, or may be learned by practice of the technology disclosed in the description.

[0011] Various turbomachinery engines and gear assemblies are disclosed herein. The disclosed turbomachinery engine includes a variable pitch fan ("VPF"). In some configurations, the disclosed engine may further include a gearbox. The disclosed variable pitch fan is characterized by a plurality of fan parameters, the plurality of fan parameters including a first VPF parameter and a second VPF parameter, the first VPF parameter being defined as the hub-to-tip radius ratio ("RR") divided by the fan pressure ratio ("FPR"), and the second VPF parameter being defined as the bearing spanwise force ("F_span") divided by the fan area ("F_area"). The disclosed VPF parameters may also be used, for example, to assist in the development of variable pitch fans and / or other engine configurations. The VPF parameters thus provide an improved variable pitch fan and / or may help simplify one or more complexities in the development of a variable pitch fan and / or geared turbomachinery engine.

[0012] In certain examples, a turbomachinery engine includes a fan assembly comprising a plurality of fan blades. The fan blades are configured such that they define a first VPF parameter and a second VPF parameter. In some examples, the first VPF parameter is in the range of 0.1 to 0.25, and the second VPF parameter is in the range of 2-30 lbf / in. 2 In other examples, the first VPF parameter is in the range of 0.1 to 0.4, and the second VPF parameter is in the range of 5.25-30 lbf / in 2In some examples, the turbomachinery engine further includes a pitch change mechanism, a blade assembly, a core engine, and a gearbox. The pitch change mechanism is coupled to a plurality of fan blades and is configured to adjust the pitch of the plurality of fan blades. The blade assembly includes a plurality of blades disposed behind the fan blades. The core engine includes one or more compressor sections and one or more turbine sections. The gearbox includes an input and an output. The input is coupled to the one or more turbine sections of the core engine and includes a first rotational speed, and the output is coupled to the fan assembly and has a second rotational speed less than the first rotational speed.

[0013] These and other features, aspects and advantages of the present disclosure will become better understood with reference to the following description and appended claims.The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate various embodiments of the disclosed technology and, together with the description, serve to explain the principles of the present disclosure. BRIEF DESCRIPTION OF THE DRAWINGS

[0014] Figure 1 Depicted is a cross-sectional schematic diagram of an exemplary embodiment of a turbomachinery engine configured with an open rotor propulsion system and a variable pitch fan.

[0015] Figure 2 Depicted is a cross-sectional schematic diagram of an exemplary embodiment of a turbomachinery engine including an open rotor propulsion system, a variable pitch fan, a three-flow architecture, and one or more heat exchangers in a third flow of the three-flow architecture.

[0016] Figure 3 Depicted is a cross-sectional schematic diagram of an exemplary embodiment of a turbomachinery engine configured with a ducted propulsion system and a variable pitch fan.

[0017] Figure 4 Depicted is a cross-sectional schematic diagram of an exemplary embodiment of a turbomachinery engine including a ducted propulsion system, a variable pitch fan, a three-flow architecture, and one or more heat exchangers in a third flow of the three-flow architecture.

[0018] Figure 5 Depicted is a schematic cross-sectional view of an exemplary embodiment of a counter-rotating low-pressure turbine of a turbomachinery engine, the low-pressure turbine having a 3x3 configuration.

[0019] Figure 6 Depicted is a schematic cross-sectional view of an exemplary embodiment of a counter-rotating low-pressure turbine of a turbomachinery engine, the low-pressure turbine having a 4x3 configuration.

[0020] Figure 7 Depicted is a cross-sectional schematic diagram of an exemplary embodiment of a gearbox configuration for a turbomachinery engine.

[0021] Figure 8 Depicted is a cross-sectional schematic diagram of an exemplary embodiment of a gearbox configuration for a turbomachinery engine.

[0022] Figure 9 Depicted is a cross-sectional schematic diagram of an exemplary embodiment of a gearbox configuration for a turbomachinery engine.

[0023] Figure 10 Depicted is a cross-sectional schematic diagram of an exemplary embodiment of a gearbox configuration for a turbomachinery engine.

[0024] Figure 11 Depicted is a cross-sectional schematic diagram of an exemplary embodiment of a gearbox configuration for a turbomachinery engine.

[0025] Figure 12 An exemplary range of a first VPF parameter is depicted relative to an exemplary range of a second VPF parameter, which may be particularly advantageous for turbomachinery including a variable pitch fan.

[0026] Figure 13 An exemplary range of a first VPF parameter is depicted relative to an exemplary range of a second VPF parameter, which may be particularly advantageous for turbomachinery including a variable pitch fan.

[0027] Figure 14 An exemplary range of a first VPF parameter is depicted relative to an exemplary range of a second VPF parameter, which may be particularly advantageous for turbomachinery including a variable pitch fan.

[0028] Figure 15 An exemplary range of a first VPF parameter is depicted relative to an exemplary range of a second VPF parameter, which may be particularly advantageous for turbomachinery including a variable pitch fan.

[0029] Figure 16 An exemplary range of a first VPF parameter is depicted relative to an exemplary range of a second VPF parameter, which may be particularly advantageous for turbomachinery including a variable pitch fan.

[0030] Figure 17 An exemplary range of a first VPF parameter is depicted relative to an exemplary range of a second VPF parameter, which may be particularly advantageous for turbomachinery including a variable pitch fan.

[0031] Figure 18 Various fan parameters are depicted for several exemplary turbomachinery engines including variable pitch fans.

[0032] Figure 19Depicted is a partial cross-sectional schematic diagram of an exemplary embodiment of a turbomachinery engine configured with a ducted propulsion system and a variable pitch fan. DETAILED DESCRIPTION

[0033] Reference will now be made in detail to embodiments of the disclosed technology, one or more examples of which are illustrated in the accompanying drawings. Each example is provided by way of explanation of the disclosed technology rather than limitation of the disclosure. In fact, it will be apparent to those skilled in the art that various modifications and variations may be made to the disclosure without departing from the scope or spirit of the disclosure. For example, a feature shown or described as part of one embodiment may be used together with another embodiment to produce yet another embodiment. Therefore, the disclosure is intended to encompass such modifications and variations as fall within the scope of the appended claims and their equivalents.

[0034] 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.

[0035] As used herein, the terms “first,” “second,” and “third” are used interchangeably to distinguish one component from another and are not intended to denote the position or importance of a single component.

[0036] The terms "fore" and "aft" refer to relative positions within a gas turbine engine or vehicle and refer to the normal operating attitude of the gas turbine engine or vehicle. For example, with respect to a gas turbine engine, the forward position refers to the position closer to the engine inlet, and the aft position refers to the position closer to the engine nozzle or exhaust.

[0037] The terms "upstream" and "downstream" refer to relative directions of fluid flow in a fluid path. For example, "upstream" refers to the direction from which the fluid is flowing, and "downstream" refers to the direction toward which the fluid is flowing.

[0038] The terms “coupled,” “fixed,” “attached,” and the like refer to both direct coupling, fixing, or attachment, as well as indirect coupling, fixing, or attachment through one or more intermediate components or features, unless otherwise specified herein.

[0039] The singular forms "a," "an," and "the" include plural references unless the context clearly dictates otherwise.

[0040] Approximating language, as used in this specification and claims, is used to modify any quantitative representation that can be permissibly varied without resulting in a change in the basic function to which it is related. Thus, a value modified by one or more terms, such as "about," "approximately," and "substantially," is not limited to the precise value specified. In at least some instances, approximating language may correspond to the precision of an instrument used to measure a value, or the precision of a method or machine used to construct or manufacture a component and / or system. For example, approximating language may refer to being within a margin of 1, 2, 4, 10, 15, or 20% of a single value, a range of values, and / or an endpoint of a defined range of values.

[0041] Here and throughout the specification and claims, range limitations are combined and interchangeable, and these ranges are identified and include all sub-ranges contained therein unless context or language indicates otherwise. For example, all ranges disclosed herein include the endpoints, and the endpoints are independently combinable with each other.

[0042] Referring now to the accompanying drawings, Figure 1 1 is an exemplary embodiment of an engine 100 including a gear assembly 102 according to aspects of the present disclosure. The engine 100 includes a fan assembly 104 driven by a core engine 106. In various embodiments, the core engine 106 is a Brayton cycle system configured to drive the fan assembly 104. The core engine 106 is at least partially enclosed by a casing 114. The fan assembly 104 includes a plurality of fan blades 108. A blade assembly 110 extends from the casing 114 in a cantilevered manner. Therefore, the blade assembly 110 may also be referred to as an unducted blade assembly. The blade assembly 110, including the plurality of blades 112, is positioned in an operative arrangement with the fan blades 108 to provide thrust, control thrust vectoring, reduce or redirect unwanted noise, and / or otherwise desirably change airflow relative to the fan blades 108.

[0043] In some embodiments, fan assembly 104 includes eight (8) to twenty (20) fan blades 108. In certain embodiments, fan assembly 104 includes ten (10) to eighteen (18) fan blades 108. In certain embodiments, fan assembly 104 includes twelve (12) to sixteen (16) fan blades 108. In some embodiments, blade assembly 110 includes three (3) to thirty (30) blades 112. In certain embodiments, blade assembly 110 includes an equal to or fewer number of blades 112 than fan blades 108. For example, in certain embodiments, engine 100 includes twelve (12) fan blades 108 and ten (10) blades 112. In other embodiments, blade assembly 110 includes a greater number of blades 112 than fan blades 108. For example, in certain embodiments, engine 100 includes ten (10) fan blades 108 and twenty-three (23) blades 112. In another particular embodiment, the engine includes fourteen (14) fan blades.

[0044] In some embodiments, for example Figure 1 As depicted in FIG, the blade assembly 110 is positioned downstream or behind the fan assembly 104. However, it should be understood that in some embodiments, the blade assembly 110 may be located upstream or ahead of the fan assembly 104. In various embodiments, the engine 100 may include a first blade assembly positioned ahead of the fan assembly 104 and a second blade assembly positioned behind the fan assembly 104. The fan assembly 104 may be configured to desirably adjust the pitch at one or more fan blades 108, such as to control thrust vectoring, reduce or redirect noise, and / or vary thrust output. The blade assembly 110 may be configured to desirably adjust the pitch at one or more blades 112, such as to control thrust vectoring, reduce or redirect noise, and / or vary thrust output. The pitch control mechanisms at one or both of the fan assembly 104 or the blade assembly 110 may operate together to produce one or more of the desired effects described above.

[0045] In some embodiments, such as Figure 1As depicted in , engine 100 is an unducted thrust generating system such that a plurality of fan blades 108 are not shielded by a nacelle or fan case. Thus, in various embodiments, engine 100 may be configured as an unshrouded turbofan engine, an open rotor engine, or a propfan engine. In a particular embodiment, engine 100 is an unducted rotor engine having a single row of fan blades 108. Fan blades 108 may have a large diameter, such as may be suitable for a high bypass ratio, a high cruise speed (e.g., compared to an aircraft having a turbofan engine, or substantially higher than the cruise speed of an aircraft having a turboprop engine), a high cruise altitude (e.g., compared to an aircraft having a turbofan engine, or substantially higher than the cruise speed of an aircraft having a turboprop engine), and / or a relatively low rotational speed.

[0046] Fan blade 108 includes a diameter D 风扇 (D fan ). It should be noted that for illustration purposes only D is shown. 风扇 In some embodiments, D 风扇 is 72-216 inches. In certain embodiments, D 风扇 is 100-200 inches. In some embodiments, D 风扇 is 120-190 inches. In other embodiments, D 风扇 In other embodiments, D 风扇 It is 50-80 inches.

[0047] In some embodiments, the fan blade tip speed at cruise flight conditions may be 650 to 900 fps, or 700 to 800 fps. The fan assembly 104 may have a fan pressure ratio (FPR) of 1.04 to 1.10, or in some embodiments, 1.05 to 1.08, as measured across the fan blades at cruise flight conditions. In other examples, the fan assembly's fan pressure ratio may be 1.05-1.5 (or 1.05-1.15 or 1.2-1.4) measured at static sea level takeoff operating conditions.

[0048] The cruising altitude is generally the altitude at which the aircraft levels off after climbing and before descending into the approach phase of flight. In various embodiments, the engine is applied to a vehicle capable of cruising altitudes up to approximately 65,000 feet. In certain embodiments, the cruising altitude is between approximately 28,000 feet and approximately 45,000 feet. In still certain embodiments, the cruising altitude is expressed in flight levels (FL) based on standard sea level pressure, wherein cruising flight conditions are between FL280 and FL650. In another embodiment, cruising flight conditions are between FL280 and FL450. In certain embodiments, the cruising altitude is defined based at least on atmospheric pressure, wherein based on a sea level pressure of approximately 14.70 psia and a sea level temperature of approximately 59 degrees Fahrenheit, the cruising altitude is between approximately 4.85 psia and approximately 0.82 psia. In another embodiment, the cruising altitude is between approximately 4.85 psia and approximately 2.14 psia. It will be appreciated that in certain embodiments, the cruising altitude range defined by pressure may be adjusted based on a different reference sea level pressure and / or sea level temperature.

[0049] The core engine 106 is generally enclosed in a casing 114 that defines a core diameter D 核心 (D core ), which can be considered as the maximum extent from the centerline axis (reference to R). In certain embodiments, the engine 100 includes a length (L) from a longitudinal (or axial) front end 116 to a longitudinal rear end 118. In various embodiments, the engine 100 defines an L / D that provides reduced installation drag. 核心 In one embodiment, L / D 核心 At least 2. In another embodiment, L / D 核心 At least 2.5. In some embodiments, L / D 核心 Less than 5, less than 4, and less than 3. In various embodiments, it should be understood that L / D 核心 For use with single unducted rotor engines.

[0050] Reduced installed drag can also provide improved efficiency, such as improved fuel consumption. Additionally or alternatively, reduced drag can provide cruising altitude engines and aircraft operating at speeds of Mach 0.5 or greater. In some embodiments, L / D 核心 , fan assembly 104 and / or blade assembly 110 individually or together at least partially configure engine 100 to operate at a maximum cruise altitude operating speed between approximately Mach 0.55 and approximately Mach 0.85 (or between approximately 0.72 to 0.85 or between approximately 0.75 to 0.85).

[0051] Still refer to Figure 1, the core engine 106 extends in a radial direction (R) relative to the engine centerline axis 120. The gear assembly 102 receives power or torque from the core engine 106 via a power input source 122 and provides power or torque in a circumferential direction C around the engine centerline axis 120 via a power output source 124 to drive the fan assembly 104.

[0052] The gear assembly 102 of the engine 100 may include a plurality of gears, including an input and an output. The gear assembly may also include one or more intermediate gears disposed between and / or interconnecting the input and output. The input may be coupled to the turbine section of the core engine 106 and may include a first rotational speed. The output may be coupled to the fan assembly and may have a second rotational speed. In some embodiments, the gear ratio of the first rotational speed to the second rotational speed is greater than 4.1 (e.g., within the range of 4.1-14.0).

[0053] The gear assembly 102 (also referred to as a "gearbox") can include various types and / or configurations. For example, in some embodiments, the gearbox is an epicyclic gearbox configured in 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 the 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 rotational speed. The planetary gears are disposed between and interconnect the sun gear and the ring gear. The planetary gears are rotatably coupled to a fixed carrier. Therefore, the planetary gears can rotate about their respective axes but cannot orbit together relative to the sun gear or the ring gear. As another example, the gearbox is an epicyclic gearbox in a planetary gear configuration. The planetary gear configuration includes a sun gear, a plurality of planetary gears, and a ring gear. The sun gear is the input and is coupled to the power turbine. The planetary gears are disposed between and interconnect the sun gear and the ring gear. The planetary gears are rotatably coupled to a rotatable carrier. Therefore, the planet gears can rotate about their respective axes and can also rotate together with the carrier relative to the sun gear and ring gear. The carrier is the output and is coupled to the fan assembly. The ring gear is fixed and does not rotate.

[0054] In some embodiments, the gearbox is a single-stage gearbox (e.g., Figure 10-11 In other embodiments, the gearbox is a multi-stage gearbox (e.g., Figure 9 and 12 In some embodiments, the gearbox is an epicyclic gearbox. In some embodiments, the gearbox is a non-episodic gearbox (e.g., a compound gearbox- Figure 13 ).

[0055] As described above, gear assembly can be used to reduce the rotational speed of output relative to input.In certain embodiments, the gear ratio of input rotational speed to output rotational speed is greater than 4.1.For example, in a specific embodiment, the gear ratio is in the range of 4.1-14.0, in the range of 4.5-14.0 or in the range of 6.0-14.0.In certain embodiments, the gear ratio is in the range of 4.5-12 or in the range of 6.0-11.0.Therefore, in certain embodiments, fan assembly can be configured to rotate with a rotational speed of 700-1500rpm under cruise flight conditions, and power turbine (e.g., low-pressure turbine) is configured to rotate with a rotational speed of 2500-15000rpm under cruise conditions.In a specific embodiment, fan assembly can be configured to rotate with a rotational speed of 850-1350rpm under cruise flight conditions, and power turbine is configured to rotate with a rotational speed of 5500-10000rpm under cruise conditions.

[0056] Various gear assembly configurations are schematically depicted in Figure 9-13 These gearboxes may be used with any engine disclosed herein, including engine 100. Additional details regarding the gearboxes are provided below.

[0057] Figure 2 A cross-sectional view of engine 200 is shown, which is configured as an exemplary embodiment of an open rotor propulsion engine. Engine 200 is generally similar to engine 100 and corresponding components have been numbered similarly. For example, the gear assembly of engine 100 is numbered "102", the gear assembly of engine 200 is numbered "202", and so on. In addition to gear assembly 202, engine 200 includes a fan assembly 204, which includes a plurality of fan blades 208 distributed about an engine centerline axis 220. Fan blades 208 are circumferentially arranged in an equally spaced relationship about engine centerline axis 220, and each fan blade 208 has a root 225 and a tip 226, and an axial span defined between root 225 and tip 226, and a center blade axis 228.

[0058] The core engine 206 includes a compressor section 230, a combustion section 232, and a turbine section 234 (which may be referred to as an "expansion section") arranged together in series flow. The core engine 206 extends circumferentially relative to the engine centerline axis 220. The core engine 206 includes a high-pressure spool including a high-pressure compressor 236 and a high-pressure turbine 238 operably and rotatably coupled together by a high-pressure shaft 240. The combustion section 232 is positioned between the high-pressure compressor 236 and the high-pressure turbine 238.

[0059] The combustion section 232 can be configured as a deflagrative combustion section, a rotating detonation combustion section, a pulse detonation combustion section, and / or other suitable heat addition systems. The combustion section 232 can be configured as one or more of a rich burn system or a lean burn system, or a combination thereof. In various embodiments, the combustion section 232 includes an annular combustor, a can combustor, a tubular combustor, a trapped vortex combustor (TVC), or other suitable combustion systems, or a combination thereof.

[0060] Core engine 206 also includes a supercharger or low-pressure compressor positioned in flow relationship with high-pressure compressor 236. Low-pressure compressor 242 is rotatably coupled to low-pressure turbine 244 via low-pressure shaft 246 to enable low-pressure turbine 244 to drive low-pressure compressor 242. Low-pressure shaft 246 is also operatively connected to gear assembly 202 to provide power to fan assembly 204, such as further described herein.

[0061] It should be understood that the terms "low" and "high," or their respective comparatives (e.g., "lower" and "higher," where applicable), when used with compressor, turbine, shaft, or spool components, refer to relative pressures and / or relative speeds within the engine, unless otherwise noted. For example, a "low spool" or "low-speed shaft" defines a component that is configured to operate at a rotational speed (such as a maximum allowable rotational speed) that is lower than the "high-speed spool" or "high-speed shaft" of the engine. Alternatively, the above terms may be understood in their superlative form, unless otherwise noted. For example, a "low turbine" or "low-speed turbine" may refer to the lowest maximum rotational speed turbine within the turbine section, a "low compressor" or "low-speed compressor" may refer to the lowest maximum rotational speed turbine within the compressor section, a "high turbine" or "high-speed turbine" may refer to the highest maximum rotational speed turbine within the turbine section, and a "high compressor" or "high-speed compressor" may refer to the highest maximum rotational speed compressor within the compressor section. Similarly, a low-speed spool refers to a lower maximum rotational speed than a high-speed spool. It should also be understood that in the above aspects, the terms "low" or "high" may additionally or alternatively be understood as relative to a minimum allowable speed, or relative to a normal, desired, steady-state, etc. minimum or maximum allowable speed for operation of the engine.

[0062] The compressors and / or turbines disclosed herein may include various numbers of stages. As disclosed herein, the number of stages includes the number of rotor or blade stages in a particular component (e.g., a compressor or turbine). For example, in some embodiments, the low-pressure compressor may include 1-8 stages, the high-pressure compressor may include 8-15 stages, the high-pressure turbine may include 1-2 stages, and / or the low-pressure turbine may include 3-7 stages. For example, in certain embodiments, the engine may include a one-stage low-pressure compressor, an 11-stage high-pressure compressor, a two-stage high-pressure compressor, and a seven-stage low-pressure turbine. As another example, the engine may include a three-stage low-pressure compressor, a 10-stage high-pressure compressor, a two-stage high-pressure compressor, and a seven-stage low-pressure turbine.

[0063] In some embodiments, the low-pressure turbine is a counter-rotating low-pressure turbine comprising inner bladed stages and outer bladed stages. The inner bladed stages extend radially outward from the inner shaft, while the outer bladed stages extend radially inward from the outer drum. In certain embodiments, the counter-rotating low-pressure turbine comprises three inner bladed stages and three outer bladed stages, which may be collectively referred to as a six-stage low-pressure turbine. In other embodiments, the counter-rotating low-pressure turbine comprises four inner bladed stages and three outer bladed stages, which may be collectively referred to as a seven-stage low-pressure turbine.

[0064] As discussed in more detail below, core engine 206 includes a gear assembly 202 configured to transfer power from turbine section 234 and reduce the output rotational speed at fan assembly 204 relative to low-pressure turbine 244. The embodiments of gear assembly 202 depicted and described herein can allow for gear ratios suitable for large diameter unducted fans (e.g., 4.1-14.0, 4.5-14.0, and / or 6.0-14.0 gear ratios). Additionally, the embodiments of gear assembly 202 provided herein can be adapted to radial or diameter constraints of core engine 206 within casing 214.

[0065] Various gearbox configurations are schematically depicted in Figure 9-13 These gearboxes may be used in any of the engines disclosed herein, including engine 200. Additional details regarding the gearboxes are provided below.

[0066] Engine 200 further includes a bucket assembly 210 comprising a plurality of buckets 212 arranged about an engine centerline axis 220. Each bucket 212 has a root 248 and a tip 250, and a span defined between the root 248 and the tip 250. The buckets 212 can be arranged in a variety of ways. For example, in some embodiments, they are not all equidistant from the rotating assembly.

[0067] In some embodiments, bucket 212 is mounted to a stationary frame and does not rotate relative to engine centerline axis 220, but may include mechanisms for adjusting its orientation relative to its axis 254 and / or relative to fan blades 208. For reference purposes, Figure 2 A forward direction is depicted, indicated by arrow F, which in turn defines the front and rear portions of the system.

[0068] like Figure 2 As depicted in FIG, the fan assembly 204 is located forward of the core engine 106, and the exhaust 256 is located aft of the core engine 206 in a "tractor" configuration. Other configurations are possible and contemplated within the scope of the present disclosure, such as embodiments that may be referred to as "pusher" configurations, in which the engine core is located forward of the fan assembly. The choice of a "tractor" or "pusher" configuration may be consistent with the choice of mounting orientation for the airframe for the intended aircraft application, and some may be structurally or operationally advantageous depending on whether the mounting location and orientation is appropriate for a wing-mounted, fuselage-mounted, or tail-mounted configuration.

[0069] Left-hand or right-hand engine configurations, which may be used in certain installations to reduce the effects of multiple engine torques on the aircraft, may be achieved by mirroring the airfoils (e.g., 208, 212) so that the fan assembly 204 rotates clockwise for one propulsion system and counterclockwise for the other propulsion system. Optionally, an optional reversing gearbox may be provided to allow the fan blades to rotate clockwise or counterclockwise using a common gas turbine core and low-pressure turbine, i.e., to provide left-hand or right-hand configurations as desired, such as to provide a pair of counter-rotating engine assemblies for certain aircraft installations without having internal engine components designed for opposite rotation directions.

[0070] Engine 200 further includes a gear assembly 202 that includes a gear set for reducing the rotational speed of fan assembly 204 relative to low-pressure turbine 244. In operation, rotating fan blades 208 are driven by low-pressure turbine 244 via gear assembly 202, causing fan blades 208 to rotate about engine centerline axis 220 and generate thrust to propel engine 200, and therefore the aircraft on which it is mounted, in a forward direction F.

[0071] In some embodiments, the gear ratio of input rotational speed to output rotational speed is greater than 4.1. For example, in a specific embodiment, the gear ratio is in the range of 4.1-14.0, in the range of 4.5-14.0, or in the range of 6.0-14.0. In certain embodiments, the gear ratio is in the range of 4.5-12 or in the range of 6.0-11.0. Therefore, in some embodiments, the fan assembly may be configured to rotate at a rotational speed of 700-1500rpm under cruise flight conditions, and the power turbine (e.g., low-pressure turbine) is configured to rotate at a rotational speed of 5000-10000rpm under cruise flight conditions. In a specific embodiment, the fan assembly may be configured to rotate at a rotational speed of 850-1350rpm under cruise flight conditions, and the power turbine is configured to rotate at a rotational speed of 5500-9500rpm under cruise flight conditions.

[0072] It may be desirable to incorporate a pitch-changing mechanism into one or both of fan blades 208 or vanes 212 so that the blades can be rotated, independently or in conjunction with one another, about a pitch axis of rotation (designated 228 and 254, respectively). Such pitch changes can be used to vary thrust and / or vortex effects under various operating conditions, including providing a thrust reversal feature that is useful under certain operating conditions, such as when landing an aircraft.

[0073] The size, shape, and configuration of the blades 212 can be arranged to impart a reactionary vortex to the fluid so that in the downstream direction behind both the fan blades 208 and the blades 212, the fluid has a significantly reduced degree of vortex, which translates into an increased level of induced efficiency. The blades 212 can have a shorter span than the fan blades 208, e.g. Figure 2 As shown. For example, the span of blade 212 can be at least 50% of the span of fan blade 208. In some embodiments, the span of blade 212 can be the same as or longer than the span of fan blade 208, if desired. Blade 212 can be attached to an aircraft structure associated with engine 200, such as Figure 2 As shown, or other aircraft structures such as wings, pylons, or fuselages. The number of blades 212 can be less than, greater than, or the same as the number of fan blades 208. In some embodiments, the number of blades 212 is greater than two or greater than four. The size, shape, and profile of fan blades 208 can take into account the expected blade loads.

[0074] exist Figure 2In the illustrated embodiment, annular 360-degree inlet 258 is located between fan assembly 204 and bucket assembly 210 and provides a path for incoming atmospheric air to pass radially inward over at least a portion of bucket assembly 210 and into core engine 206. This location may be advantageous for a variety of reasons, including managing icing performance and protecting inlet 258 from various objects and materials that may be encountered during operation.

[0075] exist Figure 2 In the exemplary embodiment, in addition to an open rotor or unducted fan assembly 204 having a plurality of fan blades 208, an optional ducted fan assembly 260 is further included behind the fan assembly 204, so that the engine 200 includes both ducted and unducted fans, both of which are used to generate thrust by moving air at ambient temperature without passing through the core engine 206. The ducted fan assembly 260 is shown at approximately the same axial position as the blades 212, and radially inward of the roots 248 of the blades 212. Alternatively, the ducted fan assembly 260 can be between the blades 212 and the core duct 262, or further forward of the blades 212. The ducted fan assembly 260 can be driven by the low-pressure turbine 244, or by any other suitable source of rotation, and can serve as the first stage of the low-pressure compressor 242 or can operate independently. Air entering the inlet 258 flows through the inlet duct 264 and is then split, with a portion flowing through the core duct 262 and a portion flowing through the fan duct 266. The fan duct 266 may incorporate one or more heat exchangers 268 and discharge to the atmosphere through a separate fixed or variable nozzle 270 located aft of the bucket assembly 210, at the aft end of the fan cowl 252, and outside the engine core cowl 272. Air flowing through the fan duct 266 thus "bypasses" the engine core and does not pass through the core.

[0076] Thus, in the exemplary embodiment, engine 200 includes an unducted fan formed by fan blades 208, followed by a ducted fan assembly 260 that directs airflow into two concentric or non-concentric ducts 262 and 266, thereby forming a three-flow engine architecture with three paths for air passing through fan assembly 204. The "first flow" of engine 200 includes airflow passing through the outside of blade assembly 210 and / or fan cowl 252. Therefore, the first flow can be referred to as a "bypass flow" because the airflow of the first flow does not pass through core duct 262. The first flow generates most of the thrust of engine 200 and, therefore, can also be referred to as the "main propulsive flow." The "second flow" of engine 200 includes airflow flowing into inlet 258, through inlet duct 264, through core duct 262, and out of core nozzle 278. In this manner, the second flow can be referred to as a "core flow." The “third flow” of engine 200 includes the airflow flowing into inlet 258 , through inlet duct 264 , through fan duct 266 , and out of nozzle 270 .

[0077] As used herein, "tertiary flow" refers to a secondary air flow that can increase the energy of the flow to produce a small portion of the total thrust of an engine (e.g., engine 200). Therefore, in various embodiments, a fan duct 266 having one or more heat exchangers 268 located within the flow path of the fan duct 266 can be referred to as the "tertiary flow" of a three-flow engine architecture.

[0078] The pressure ratio of the tertiary flow is higher than the pressure ratio of the main propulsive flow (i.e., the bypass flow). This thrust is generated through a dedicated nozzle or by mixing the secondary flow with the fan flow or core flow (e.g., entering a common nozzle). In certain exemplary embodiments, the operating temperature of the airflow through the tertiary flow is lower than the maximum compressor discharge temperature of the engine, and more specifically can be lower than 350 degrees Fahrenheit (e.g., lower than 300 degrees Fahrenheit, such as lower than 250 degrees Fahrenheit, such as lower than 200 degrees Fahrenheit, and at least as high as the ambient temperature). In certain exemplary embodiments, these operating temperatures are conducive to heat transfer to or from the fluids in the tertiary flow and the secondary fluid flow. Further, in certain exemplary embodiments, under takeoff conditions, or more specifically, when operating at sea level at rated takeoff power, static flight speed, and 86 degrees Fahrenheit ambient temperature operating conditions, the airflow through the tertiary flow can contribute less than 50% of the total engine thrust (and at least, for example, 2% of the total engine thrust). Furthermore, in certain exemplary embodiments, the airflow, mixing, or exhaust characteristics of the third stream (and thereby the aforementioned exemplary percentage contribution to total thrust) can be passively adjusted during engine operation or purposefully modified through the use of engine control features (such as fuel flow, motor power, variable stators, variable inlet guide vanes, valves, variable exhaust geometry, or fluid ground features) to adjust or optimize overall system performance over a wide range of potential operating conditions.

[0079] exist Figure 2In the exemplary embodiment shown, a slidable, movable, and / or translatable plug nozzle 274 with an actuator may be included to vary the outlet area of ​​the nozzle 270. A plug nozzle is typically an annularly symmetrical device that adjusts the opening area of ​​an outlet, such as fan flow or core flow, by axially moving the nozzle so that the gap between the nozzle surface and a stationary structure (e.g., an adjacent wall of a duct) varies in a predetermined manner, thereby reducing or increasing the space for airflow through the duct. Other suitable nozzle designs may also be employed, including those incorporating thrust reversal functionality. Such adjustable, movable nozzles may be designed to operate in conjunction with other systems (e.g., VBV, VSV, or blade pitch mechanisms) and may be designed with failure modes (e.g., fully open, fully closed, or an intermediate position) such that the nozzle 270 has a consistent "home" position to which it returns in the event of any system failure, which may prevent commands from reaching the nozzle 270 and / or its actuator. In other embodiments, a static nozzle may be used.

[0080] In some embodiments, a mixing device 276 may be included in the region behind the core nozzle 278 to help mix the fan flow and the core flow, thereby improving acoustic performance by directing the core flow outward and the fan flow inward.

[0081] because Figure 2 The illustrated engine 200 includes an open rotor fan assembly 204, a ducted fan assembly 260, and a third flow. The engine's thrust output and work distribution can be tailored to achieve specific thrust, fuel burn, thermal management, and / or acoustic signature targets, potentially outperforming typical ducted or unducted fan gas turbine propulsion assemblies of comparable thrust ratings. Operationally, engine 200 can include a control system that manages the loads of the corresponding open and ducted fans, as well as the exit area of ​​the potential variable fan nozzle, to provide different thrust, noise, cooling capacity, and other performance characteristics for various portions of the flight envelope and various operating conditions associated with aircraft operations. For example, in climb mode, the ducted fan can be operated at a maximum pressure ratio to maximize the thrust capability of the airflow, while in cruise mode, the ducted fan can be operated at a lower pressure ratio to improve overall efficiency by relying on the thrust of the unducted fan. Nozzle actuation adjusts the ducted fan operating line and overall engine fan pressure ratio independently of the total engine airflow. In other embodiments, static nozzles can be used to manage loading.

[0082] As described above, the third flow (e.g., fan duct 266) may include one or more heat exchangers 268 to remove heat from various fluids used in engine operation, such as an air-cooled oil cooler (ACOC), cooled cooling air (CCA), etc. Compared to conventional ducted fan architectures, heat exchangers 268 located in the third flow take advantage of being integrated into the fan duct 266, reducing performance penalties (e.g., fuel efficiency and thrust) by not impacting the primary source of thrust, in this case, the unducted fan flow. Heat exchangers can cool fluids (e.g., gearbox oil, engine sump oil), heat transfer fluids (e.g., supercritical fluids), or commercially available single- or two-phase fluids (supercritical CO2, EGV, Slither 800, liquid metal, etc.), engine exhaust, etc. Heat exchangers can also be composed of different segments or channels cooling different working fluids, such as an ACOC paired with a fuel cooler. Heat exchangers 268 can be incorporated into a thermal management system that provides heat transfer via a heat exchange fluid flowing through a network to remove heat from a source and transfer it to the heat exchanger.

[0083] Because the fan pressure ratio associated with the tertiary flow is higher for ducted fans than for unducted fans, the tertiary flow provides an environment in which a more compact heat exchanger can be used than would be possible if installed outside the core shroud in an unducted fan flow. The fan pressure ratio (FPR) of the fan bypass air is very low (1.05 to 1.08), making it difficult to drive air through the heat exchanger. In the absence of fan ducting as described herein, it may be necessary to collect or pressurize the exhaust gas to provide cooling air to and through the heat exchanger.

[0084] The third flow can also provide other advantages in terms of reducing nacelle drag, achieving more positive nacelle closure, improved core flow particle separation, and bad weather operation. By discharging the fan duct flow at the core cowl, this helps to stimulate the boundary layer and enables the selection of a steeper nacelle closure angle between the maximum size of the engine core cowl 272 and the exhaust port 256. The closure angle is usually limited by airflow separation, but the boundary layer stimulated by the air from the fan duct 266 discharged at the core cowl reduces airflow separation. This results in a shorter, lighter structure with less friction surface drag.

[0085] The fan assembly and / or blade assembly may be shielded or unshielded (e.g. Figure 1 and 2Although not shown, an optional annular shroud or duct may be coupled to the bucket assembly 210 and located distally of the engine centerline axis 220 relative to the buckets 212. In addition to the noise reduction benefits, the duct may provide improved vibration response and structural integrity of the buckets 212 by coupling it to the assembly forming an annular ring or one or more circumferential sectors (i.e., segments forming portions of an annular ring connecting two or more buckets 212). The duct may also allow the pitch of the buckets to be more easily varied. For example, as discussed in more detail below, Figure 3-4 Embodiments are disclosed in which both the fan assembly and the bucket assembly are shielded.

[0086] Although described in the above embodiments as unshrouded or open rotor engines, it should be understood that aspects of the disclosure provided herein may be applied to shrouded or ducted engines, partially ducted engines, rear fan engines, or other turbomachinery configurations, including those used in marine, industrial, or aerospace propulsion systems. Certain aspects of the disclosure may be applicable to turbofan engines, turboprop engines, or turboshaft engines. However, it should be understood that certain aspects of the disclosure may address issues that may be particularly specific to unshrouded or open rotor engines, such as, but not limited to, engine gear ratio, fan diameter, fan speed, length (L), engine core maximum diameter (D), and engine speed. 核心 ), engine L / D 核心 , desired cruise altitude, and / or desired operational cruise speed, or a combination thereof.

[0087] Unducted engines 100 and 200 may include a pitch change mechanism configured to adjust the pitch of the fan. In this manner, the fan of engines 100 and 200 is a VPF. For example, engine 200 includes a pitch change mechanism 282 coupled to fan assembly 204 and configured to change the pitch of fan blades 208. In certain embodiments, pitch change mechanism 282 may be a linear drive pitch change mechanism.

[0088] Figure 3 is a schematic cross-sectional view of a gas turbine engine according to an exemplary embodiment of the present disclosure. More specifically, Figure 3 In the embodiment of the present invention, the gas turbine engine is a high bypass turbofan jet engine 300, referred to herein as "turbofan engine 300". Figure 3As shown, the turbofan engine 300 defines an axial direction A (extending parallel to a longitudinal centerline 302 provided as a reference) and a radial direction R (extending perpendicular to the axial direction A). Generally, the engine 300 includes a fan section 304 and a core engine 306 disposed downstream of the fan section 304. The engine 300 further includes a gear assembly or power gearbox 336 having a plurality of gears for coupling the gas turbine shaft to the fan shaft. The location of the power gearbox 336 is not limited to the location shown in the exemplary embodiment of the engine 300. For example, the location of the power gearbox 336 can vary along the axial direction A.

[0089] The exemplary core engine 306 shown generally includes a substantially tubular casing 308 defining an annular inlet 310. Casing 308 surrounds, in series flow relationship, a compressor section including a supercharger or low-pressure (LP) compressor 312 and a high-pressure (HP) compressor 314; a combustion section 316; a turbine section including a high-pressure (HP) turbine 318 and a low-pressure (LP) turbine 320; and an exhaust nozzle section 322. A high-pressure (HP) shaft or spool 324 drivingly connects the HP turbine 318 to the HP compressor 314. A low-pressure (LP) shaft or spool 326 drivingly connects the LP turbine 320 to the LP compressor 312. Additionally, the compressor section, combustion section 316, and turbine section together at least partially define a core air flow path 327 extending therethrough.

[0090] The gear assembly of the present disclosure is compatible with standard fans, variable pitch fans, or other configurations. For the depicted embodiment, the fan section 304 includes a variable pitch fan 328 having a plurality of fan blades 330 coupled to a disk 332 in a spaced-apart manner. As shown, the fan blades 330 extend outwardly from the disk 332 generally in a radial direction R. Each fan blade 330 is rotatable relative to the disk 332 about a pitch axis P because the fan blades 330 are operably coupled to a suitable drive member 334, which is configured to collectively change the pitch of the fan blades 330. The fan blades 330, disk 332, and drive member 334 are rotatable together about the longitudinal axis 302 via the LP shaft 326 over a gear assembly or power gearbox 336. The gear assembly 336 can achieve speed variation between a first shaft (e.g., the LP shaft 326) and a second shaft (e.g., the LP compressor shaft and / or the fan shaft). For example, in one embodiment, the gear assembly 336 may be disposed in an arrangement between a first shaft and a second shaft to reduce the output speed from one shaft to the other shaft.

[0091] More generally, the gear assembly 336 can be placed anywhere along the axial direction A to separate the speeds of the two shafts wherever it is convenient to do so from a component efficiency standpoint, e.g., a faster LP turbine and slower fan and LP compressor or a faster LP turbine and LP compressor and a slower fan.

[0092] Still refer to Figure 3 In the exemplary embodiment, disk 332 is covered by a rotatable forward nacelle 338 having an aerodynamic profile to facilitate airflow through the plurality of fan blades 330. Furthermore, exemplary fan section 304 includes an annular fan case or outer nacelle 340 that circumferentially surrounds fan 328 and / or at least a portion of core engine 306. For the depicted embodiment, nacelle 340 is supported relative to core engine 306 by a plurality of circumferentially spaced outlet guide vanes 342. Additionally, a downstream section 344 of nacelle 340 extends over the exterior of core engine 306 to define a bypass airflow passage 346 therebetween.

[0093] During operation of turbofan engine 300, a volume of air 348 enters engine 300 through nacelle 340 and / or associated inlet 350 of fan section 304. As volume of air 348 passes through fan blades 330, a first portion of air 348, indicated by arrow 352, is directed or routed into bypass airflow passage 346, and a second portion of air 348, indicated by arrow 354, is directed or routed into LP compressor 312. The ratio between first portion of air 352 and second portion of air 354 is generally referred to as a bypass ratio. The pressure of second portion of air 354 then increases as it passes through high pressure (HP) compressor 314 and enters combustion section 316, where it mixes with fuel and combusts to provide combustion gases 356.

[0094] The combustion gases 356 are routed through the HP turbine 318, where a portion of the thermal and / or kinetic energy from the combustion gases 356 is extracted via successive stages of HP turbine stator blades 358 coupled to the casing 308 and HP turbine rotor blades 360 coupled to the HP shaft or spool 324, thereby causing the HP shaft or spool 324 to rotate, thereby supporting operation of the HP compressor 314. The combustion gases 356 are then routed through the LP turbine 320, where a second portion of the thermal and kinetic energy is extracted from the combustion gases 356 via successive stages of LP turbine stator blades 362 coupled to the casing 308 and LP turbine rotor blades 364 coupled to the LP shaft or spool 326, thereby causing the LP shaft or spool 326 to rotate, thereby supporting operation of the LP compressor 312 and / or rotation of the fan 328.

[0095] The combustion gases 356 are then routed through the jet exhaust nozzle section 322 of the core engine 306 to provide propulsive thrust. Simultaneously, the pressure of the first portion of air 352 is significantly increased as the first portion of air 352 is routed through the bypass airflow passage 346 before being discharged from the fan nozzle exhaust section 366 of the engine 300, also providing propulsive thrust. The HP turbine 318, the LP turbine 320, and the jet exhaust nozzle section 322 at least partially define a hot gas path 368 for routing the combustion gases 356 through the core engine 306.

[0096] For example, Figure 4 is a cross-sectional schematic diagram of an exemplary embodiment of an engine 400, which includes a gear assembly 402 in combination with a ducted fan assembly 404 and a core engine 406. However, unlike the open rotor configuration of engine 200, fan assembly 404 and its fan blades 408 are contained within an annular fan case 480 (also referred to as a "nacelle"), and blade assembly 410 and blades 412 extend radially between fan cowl 452 (and / or engine core cowl 472) and the inner surface of fan case 480, thereby defining a bypass flow. As described above, the gear assembly disclosed herein can provide an increased gear ratio for a fixed gear envelope (e.g., with the same size ring gear), or alternatively, a smaller diameter ring gear can be used to achieve the same gear ratio.

[0097] Core engine 406 includes a compressor section 430, a combustor section 432, and a turbine section 434. Compressor section 430 may include a high-pressure compressor 436 and a supercharger or low-pressure compressor 442. Turbine section 434 may include a high-pressure turbine 438 and a low-pressure turbine 444. Low-pressure compressor 442 is positioned forward of high-pressure compressor 436 and is in flow relationship with high-pressure compressor 436. Low-pressure compressor 442 is rotatably coupled to low-pressure turbine 444 via low-pressure shaft 446, enabling low-pressure turbine 444 to drive low-pressure compressor 442 (and ducted fan 460). Low-pressure shaft 446 is also operably connected to gear assembly 402 to provide power to fan assembly 404. High-pressure compressor 436 is rotatably coupled to high-pressure turbine 438 via high-pressure shaft 440, enabling high-pressure turbine 438 to drive high-pressure compressor 436.

[0098] A portion of the airflow from ducted fan 460 may be directed into core engine 406 (i.e., a second flow). Another portion of the airflow from ducted fan 460 may be directed into third flow 407 defined by the inner surface of fan cowling 452 and the outer surface of engine core cowling 472. In some examples, the third flow may include one or more heat exchangers.

[0099] Engine 400 (and / or engine 300) includes a pitch change mechanism 482 coupled to fan assembly 404 and configured to change the pitch of fan blades 408. In certain embodiments, pitch change mechanism 482 may be a linear drive pitch change mechanism.

[0100] In some embodiments, engine 400 may include a variable fan nozzle. Operationally, engine 400 may include a control system that manages fan loads and, possibly, the exit area of ​​the variable fan nozzle to provide different thrust, noise, cooling capacity, and other performance characteristics for various portions of the flight envelope and various operating conditions associated with aircraft operation. For example, nozzle actuation adjusts the fan operating line and overall engine fan pressure ratio independently of total engine airflow. In other examples, the engine may include a static nozzle.

[0101] In some embodiments, an engine (eg, engine 100, engine 200, and / or engine 400) may include a counter-rotating low-pressure turbine. Figure 5-6 A schematic cross-sectional view of a counter-rotating low-pressure turbine is depicted. In particular, Figure 5 A counter-rotating turbine 500 is depicted, and Figure 6 Depicted is a counter-rotating turbine 600. The counter-rotating turbine comprises inner and outer bladed stages arranged in an alternating inner and outer configuration. In other words, the counter-rotating turbine does not include stator vanes disposed between the bladed stages.

[0102] refer to Figure 5 , counter-rotating turbine 500 includes a plurality of inner blade stages 502 and a plurality of outer blade stages 504. More specifically, counter-rotating turbine 500 includes three inner blade stages 502 coupled to and extending radially outward from an inner shaft 506 (also referred to as a "rotor") and three outer blade stages 504 coupled to and extending radially inward from an outer shaft 508 (also referred to as a "drum"). In this manner, counter-rotating turbine 500 can be considered a six-stage turbine.

[0103] refer to Figure 6 , counter-rotating turbine 600 includes a plurality of inner blade stages 602 and a plurality of outer blade stages 604. More specifically, counter-rotating turbine 600 includes four inner blade stages 602 coupled to and extending radially outward from inner shaft 606 and three outer blade stages 604 coupled to and extending radially inward from outer shaft 608. In this manner, counter-rotating turbine 600 can be considered a seven-stage turbine.

[0104] According to some embodiments, there are turbomachines characterized by high gear ratios. A high gear ratio gearbox refers to a gearbox having a gear ratio greater than about 4:1 to about 14:1 (or about 4.5:1 to about 12:1 in certain embodiments). For example, the engines disclosed herein may include a gearbox configured such that the output speed (i.e., the speed of the propeller) is about 400-1200 rpm under cruise flight conditions, or more specifically, 450-1000 rpm under cruise flight conditions.

[0105] Various exemplary gear assemblies are shown and described herein. In particular, Figure 7-11 Several exemplary gear assemblies that can be used with engines 100, 200, 300, and 400 are schematically depicted. The disclosed gear assemblies can be used with any of the exemplary engines and / or any other suitable engine that may require such a gear assembly. In this manner, it should be understood that the gear assemblies disclosed herein can generally operate with an engine having a rotating element with a plurality of rotor blades and a turbomachinery having a turbine and a shaft rotatable with the turbine. For such an engine, a rotating element (e.g., fan assembly 104) can be driven by a shaft (e.g., a low-pressure shaft) of the turbomachinery via the gear assembly.

[0106] Although the exemplary gear assembly shown is mounted in a forward position (e.g., forward from the combustor and / or low-pressure compressor), in other embodiments, the gear assembly described herein may be mounted in an aft position (e.g., aft of the combustor and / or low-pressure turbine).

[0107] Various embodiments of the gear assemblies provided herein can allow for gear ratios up to 14: 1. Various embodiments of the gear assemblies provided herein can allow for gear ratios of at least 4.1: 1 or 4.5: 1. Still other various embodiments of the gear assemblies provided herein can allow for gear ratios of 6: 1 to 12: 1.

[0108] Figure 7 Schematically depicted is a gearbox 700 that may be used, for example, with engines 100, 200, 300, 400. The gearbox 700 comprises a two-stage star configuration.

[0109] The first stage of gearbox 700 includes a first-stage sun gear 702, a first-stage carrier 704 housing a plurality of first-stage planetary gears, and a first-stage ring gear 706. First-stage sun gear 702 can be coupled to a low-pressure shaft 708, which in turn is coupled to the engine's low-pressure turbine. First-stage sun gear 702 can mesh with the first-stage planetary gears, which in turn mesh with the first-stage ring gear. First-stage carrier 704 can be fixed to prevent rotation by a support member 710.

[0110] The second stage of gearbox 700 includes a second-stage sun gear 712, a second-stage carrier 714 that houses a plurality of second-stage planetary gears, and a second-stage ring gear 716. Second-stage sun gear 712 can be coupled to shaft 718, which in turn is coupled to first-stage ring gear 706. Second-stage carrier 714 can be fixed from rotation by support members 720. Second-stage ring gear 716 can be coupled to fan shaft 722.

[0111] In some embodiments, each stage of gearbox 700 may include five spider gears. In other embodiments, gearbox 700 may include fewer or more than five spider gears in each stage. In some embodiments, the first-stage carrier may include a different number of spider gears than the second-stage carrier. For example, the first-stage carrier may include five spider gears and the second-stage carrier may include three spider gears, or vice versa.

[0112] In some embodiments, the radius R1 of the gearbox 700 may be approximately 16-19 inches. In other embodiments, the radius R1 of the gearbox 700 may be approximately 22-24 inches. In other embodiments, the radius R1 of the gearbox 700 may be less than 16 inches or greater than 24 inches.

[0113] Figure 8 A gearbox 800 is schematically depicted that can be used, for example, in engines 100, 200, 300, and 400. Gearbox 800 includes a single-stage planetary configuration. Gearbox 800 includes a sun gear 802, a carrier 804 that accommodates a plurality of planetary gears (e.g., 3-5 planetary gears), and a ring gear 806. Sun gear 802 can mesh with the planetary gears, and the planetary gears can mesh with the ring gear 806. Sun gear 802 can be coupled to a low-pressure shaft 808, which in turn is coupled to a low-pressure turbine of the engine. Carrier 804 can be fixed from rotation by a support member 810. Ring gear 806 can be coupled to a fan shaft 812.

[0114] In some embodiments, the radius R2 of the gearbox 800 can be approximately 18-23 inches. In other embodiments, the radius R2 of the gearbox 700 can be less than 18 inches or greater than 23 inches.

[0115] Figure 9A gearbox 900 is schematically depicted that can be used, for example, in engines 100, 200, 300, and 400. Gearbox 900 includes a single-stage planetary configuration. Gearbox 900 includes a sun gear 902, a carrier 904 that accommodates a plurality of planetary gears (e.g., 3-5 planetary gears), and a ring gear 906. Sun gear 902 can mesh with the planetary gears, and the planetary gears can mesh with the ring gear 906. Sun gear 902 can be coupled to a low-pressure shaft 908, which in turn is coupled to a low-pressure turbine of the engine. Carrier 904 can be fixed from rotation by a support member 910. Ring gear 906 can be coupled to a fan shaft 912.

[0116] In some embodiments, the radius R3 of the gearbox 900 can be approximately 10-13 inches. In other embodiments, the radius R3 of the gearbox 900 can be less than 10 inches or greater than 13 inches.

[0117] Figure 10 Schematically depicted is a gearbox 1000 that may be used, for example, in engines 100, 200, 300, 400. The gearbox 1000 comprises a two-stage configuration, wherein the first stage is a star configuration and the second stage is a planetary configuration.

[0118] The first stage of gearbox 1000 includes a first-stage sun gear 1002, a first-stage spider carrier 1004 including a plurality of first-stage spider gears (e.g., 3-5 spider gears), and a first-stage ring gear 1006. First-stage sun gear 1002 can mesh with the first-stage spider gears, which can mesh with first-stage ring gear 1006. First-stage sun gear 1002 can be coupled to a higher-speed shaft 1008 of a low-spool, which in turn is coupled to the inner blades of the engine's low-pressure turbine. First-stage spider carrier 1004 can be fixed from rotation by a support member 1010.

[0119] The second stage of gearbox 1000 includes a second-stage sun gear 1012, a second-stage planet carrier 1014 including a plurality of second-stage planet gears (e.g., 3-5 planet gears), and a second-stage ring gear 1016. Second-stage sun gear 1012 can mesh with the second-stage planet gears. Second-stage planet carrier 1014 can be coupled to first-stage ring gear 1006. Second-stage sun gear 1012 can be coupled to a lower-speed shaft 1018 of a low-spool, which in turn is coupled to the outer blades of the low-pressure turbine of engine 4. Second-stage planet carrier 1014 can be coupled to first-stage ring gear 1006. Second-stage planet carrier 1014 can also be coupled to fan shaft 1020. Second-stage ring gear 1016 can be fixed to rotation by support member 1022.

[0120] In some embodiments, each stage of gearbox 1000 may include three planetary / star gears. In other embodiments, gearbox 1000 may include fewer or more than three planetary / star gears in each stage. In some embodiments, the first-stage carrier may include a different number of planetary gears than the second-stage carrier, which has planetary gears. For example, the first-stage carrier may include five planetary gears, and the second-stage carrier may include three planetary gears, or vice versa.

[0121] Because the first stage of gearbox 1000 is coupled to the higher-speed shaft 1008 of the low spool, and the second stage of gearbox 1000 is coupled to the lower-speed shaft 1018 of the low spool, the gear ratio of the first stage of gearbox 1000 can be greater than the gear ratio of the second stage of the gearbox. For example, in certain embodiments, the first stage of the gearbox can include a gear ratio of 4.1-14, and the second stage of the gearbox can include a gear ratio that is less than the gear ratio of the first stage of the gearbox. In a particular embodiment, the first stage of the gearbox can include a gear ratio of 7, and the second stage of the gearbox can include a gear ratio of 6.

[0122] In some embodiments, the engine including the gearbox 1000 can be configured such that the higher-speed shaft 1008 provides approximately 50% of the power to the gearbox 1000 and the lower-speed shaft 1018 provides approximately 50% of the power to the gearbox 1000. In other embodiments, the engine including the gearbox 1000 can be configured such that the higher-speed shaft 1008 provides approximately 60% of the power to the gearbox 1000 and the lower-speed shaft 1018 provides approximately 40% of the power to the gearbox 1000.

[0123] In some embodiments, the radius R4 of the gearbox 1000 can be approximately 18-22 inches. In other embodiments, the radius R4 of the gearbox 700 can be less than 18 inches or greater than 22 inches.

[0124] Figure 11 A gearbox 1100 is depicted that can be used, for example, in engines disclosed herein (e.g., engines 100, 200, 300, 400). Gearbox 1100 is configured as a compound planetary gearbox. Gearbox 1100 includes a sun gear 1102 and a planetary carrier 1104, which includes a plurality of compound planetary gears having one or more first portions 1106 and one or more second portions 1108. Gearbox 1100 further includes a ring gear 1110. Sun gear 1102 can also mesh with the first portion 1106 of the planetary gears. The planetary carrier can be fixed against rotation by a support member 1114. The second portion 1108 of the planetary gears can mesh with the ring gear 1110. Sun gear 1102 can be coupled to a low-pressure turbine via a turbine shaft 1112. Ring gear 1110 can be coupled to a fan shaft 1116.

[0125] The gear assembly shown and described herein may be used with any suitable engine. For example, while Figure 4 Shown are optional ducted fans and optional fan ducts (similar to Figure 2 ), it will be appreciated that such a gear assembly may be used with other ducted turbofan engines (e.g., engine 300) and / or other open rotor engines that do not have one or more such structures.

[0126] Figure 1-4 The depicted engine is configured such that the fan assembly and the core engine are concentric (i.e., rotate about a common axis, which may also be referred to as "on-axis"). In other embodiments, the engine may be configured such that the fan assembly rotates about a first axis and the core engine rotates about a second, non-concentric (also referred to as "eccentric") axis.

[0127] Each embodiment of the turbomachinery disclosed herein includes a variable pitch fan ("VPF") and a drive member (e.g., drive member 334), which may also be referred to as a pitch changing mechanism. The disclosed engine may also include a gearbox. Employing a variable pitch fan provides one or more advantages (e.g., increased propulsion efficiency) and also presents significant challenges. For example, incorporating variable pitch fan blades presents challenges to mechanical packaging and mechanical integration, including the packaging and integration of the drive member and the coupling between the blades and the drive member. In contrast, a turbomachinery having fan blades with a fixed pitch is relatively simpler to implement. For example, when a fixed pitch turbomachinery is employed, a reduction in the fan radius ratio (defined below) is more easily achieved because less space is required to package and integrate the fan blades below the blade root, for example, because the blades are attached directly to the fan disk rather than through a bearing assembly. The desire to provide variable pitch capability for the fan blades while achieving an acceptable fan radius ratio, and to achieve a turbomachinery design incorporating such capability while meeting other necessary requirements such as acceptable reliability for its intended use, mission requirements, etc., safety margins in the event of debris impact and damage to the fan blades, and accessibility for maintenance, blade replacement, etc. of the variable pitch fan, presents significant challenges that need to be overcome.

[0128] Based on this, the inventors set out to define the various requirements for a variable pitch fan and then constructed various embodiments to meet these varying requirements. During the development of the aforementioned embodiments of a turbomachine engine incorporating a variable pitch fan, the inventors unexpectedly discovered that a number of specific fan parameters, arranged in unique combinations, provided a good approximation for the overall variable pitch fan design. More specifically, the inventors discovered that certain value ranges defining embodiments of a variable pitch fan, including but not limited to bearing size, shape, orientation, materials, and the like, could inform a skilled artisan about the positive and negative attributes of selecting one embodiment over another, as a function of turbomachinery performance requirements. Thus, the inventors realized they had discovered the value associated with selecting one design over another based on engine requirements (e.g., blade size, fan tip speed, packaging, integration, etc.), defining not only the benefits but also the penalties associated with such selection. The embodiments defined by these VPF parameters, as they are called, therefore offer significant benefits because they limit the design space to a reduced number of practical embodiments based on the basic structural requirements needed to meet the requirements. The structural requirements involved are related to requirements such as achieving a specific weight, size, drag, and / or other factors related to the mechanical packaging of the VPF. For example, the VPF parameter ranges disclosed herein take into account limitations (eg, bearing stresses) and thus allow for adequate mechanical integration. One particularly advantageous aspect discovered by the inventors is that the VPF parameters can be used for either unducted or ducted fan designs.

[0129] The inventors have discovered that two VPF parameters are particularly important. The first VPF parameter is defined as the fan's hub-to-tip radius ratio ("RR") divided by the fan pressure ratio ("FPR") measured under static sea-level takeoff operating conditions. The second VPF parameter is defined as the fan's bearing spanwise force ("F_span") (measured in pounds-force under redline operating conditions) divided by the fan area ("F_area") (measured in square inches).

[0130] As used herein, fan radius ratio is defined as the fan hub radius (R_hub) divided by the fan tip radius (R_tip), both measured at the leading edge of the fan blades. Figure 19 , including exemplary fans of various sizes, are depicted in FIG. In some examples, the fan radius ratio is in the range of 0.125-0.55. In other examples, the fan radius ratio is in the range of 0.2-0.5. In a specific example, the fan radius ratio is in the range of 0.25-0.35.

[0131] The fan pressure ratio is defined as the ratio of the total pressure across the fan (outlet / inlet) during static sea level takeoff (SLTO) operating conditions. In some examples, the fan pressure ratio during SLTO operating conditions is in the range of 1.05-1.5. In other examples, the fan pressure ratio during SLTO operating conditions is in the range of 1.05-1.15, which (in some cases) may correspond to an unducted fan. In a specific example, the fan pressure ratio during SLTO operating conditions is in the range of 1.2-1.4, which (in some cases) may correspond to a ducted fan.

[0132] The fan blade bearing span force (F_span) is defined as (mass of the fan blade / 386.4)*R_cg*ω 2 , where R_cg is the radius of the fan blade's center of gravity measured from the fan's axis of rotation (in inches), and ω is the fan's redline speed measured in radians per second. The center of gravity, and therefore R_cg, can be calculated or approximated in various ways. As an example, R_cg can be approximated by the following equation: R_hub + 1 / 3 * (R_tip - R_hub). F_span is measured in pounds-force (lbf). In some examples, F_span is in the range of 20,000-200,000 lbf under redline operating conditions. In other examples, F_span is in the range of 50,000-100,000 lbf under redline operating conditions.

[0133] The fan area (F_area) is equal to π*(R_tip 2 -R_hub 2 ), which results in an area in square inches. In some examples, F_area is between 3,000-25,000in 2 In other examples, F_area is between 5,000-15,000in. 2 In this specific example, F_area is between 4,000-8,000in. 2 within the range.

[0134] Now refer to Figure 12 In some embodiments, the VPF may be configured such that a first VPF parameter (ie, RR / FPR) is or is approximately 0.1-0.25 and a second VPF parameter (ie, F_span / Fan_area) is or is approximately 2-30 lbf / in 2 , or the first VPF parameter is or is about 0.1-0.4 and the second VPF factor is or is about 5.25-30 lbf / in 2 .

[0135] like Figure 13 As shown, in certain embodiments, the VPF may be configured such that a first VPF parameter is or is approximately 0.1-0.25 and a second VPF parameter is or is approximately 2-30 lbf / in 2 .

[0136] refer to Figure 14 In some embodiments, the VPF may be configured such that a first VPF parameter is or is approximately 0.1-0.4 and a second VPF parameter is or is approximately 5.25-30 lbf / in 2 For example, this VPF parameter range is particularly beneficial for ducted fans.

[0137] like Figure 15 As shown, in some embodiments, the VPF may be configured such that a first VPF parameter is or is approximately 0.1-0.25 and a second VPF parameter (ie, F_span / Fan_area) is or is approximately 2-30 lbf / in 2 and the first VPF parameter is or is about 0.1-0.4 and the second VPF factor is or is about 5.25-30 lbf / in 2 In other words, the VPF may be configured such that the first VPF parameter is or is approximately 0.1-0.25 and the second VPF parameter is or is approximately 5.25-30 lbf / in 2 .

[0138] refer to Figure 16 In some embodiments, the VPF may be configured such that a first VPF parameter is or is approximately 0.25-0.4 and a second VPF parameter is or is approximately 5.25-30 lbf / in 2 .

[0139] refer to Figure 17 In certain embodiments, the VPF may be configured such that a first VPF parameter is or is approximately 0.1-0.25 and a second VPF parameter is or is approximately 5.25-30 lbf / in 2 For example, this VPF parameter range is particularly beneficial for unducted fans.

[0140] Figure 18 Depicts including falling Figure 12-17 Multiple engines (unducted and ducted) with VPFs within one or more disclosed ranges of the VPF parameters depicted. Figure 18 Various other parameters of the VPF are also plotted. Figure 18 , the fan terminal speed, fan redline speed (ω), and fan bearing span force (F_span) are listed at redline operating conditions, and the fan pressure ratio (FPR) is listed at static sea level takeoff operating conditions.

[0141] therefore, Figure 18 The exemplary engines listed in define variable pitch fans compatible with achieving, for example, specific weight, size, and / or drag requirements and / or other factors related to the mechanical packaging of the engine. As previously mentioned, the disclosed engine and its VPF take into account other constraints (e.g., bearing stresses) to allow for adequate mechanical integration. Figure 18 As shown, the disclosed VPF parameters are applicable to both unducted and ducted VPF designs. Figure 18 As further shown in the , the relationship between VPF parameters and turbomachinery embodiments discovered by the inventors provides a relatively quick and direct way to determine the feasibility of a particular VPF design. Thus, the methods and VPF parameters disclosed by the inventors can improve the design of ducted and unducted turbomachinery engines.

[0142] This written description uses examples to disclose the technology, including the best mode, and also to enable those skilled in the art to practice the technology of the present disclosure, including making and using any devices or systems and performing any combined methods. The patentable scope of the disclosed technology 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 have 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 language of the claims.

[0143] Further aspects of the present disclosure are provided by the subject matter of the following examples:

[0144] Example 1. A turbomachinery engine comprising a fan assembly, a pitch change mechanism, a blade assembly, a core engine, and a gearbox. The fan assembly comprises a plurality of fan blades, a first VPF parameter, and a second VPF parameter. The first VPF parameter is defined as the fan blade radius ratio (RR) divided by the fan pressure ratio (FPR) at static sea level takeoff operating conditions. The second VPF parameter is defined as the bearing span force (F_Span) measured in pounds-force divided by the fan area (F_Area) measured in square inches at redline operating conditions. The first VPF parameter is in the range of 0.1 to 0.25 and the second VPF parameter is in the range of 2-30 lbf / in 2 or the first VPF parameter is in the range of 0.1 to 0.4 and the second VPF parameter is in the range of 5.25-30 lbf / in 2The pitch changing mechanism is coupled to the plurality of fan blades and is configured to adjust the pitch of the plurality of fan blades. The blade assembly includes a plurality of blades disposed behind the fan blades. The core engine includes one or more compressor sections and one or more turbine sections. The gearbox includes an input and an output. The input is coupled to the one or more turbine sections of the core engine and includes a first rotational speed, and the output is coupled to the fan assembly and has a second rotational speed less than the first rotational speed.

[0145] Example 2. The turbomachinery engine according to any example herein, in particular Example 1, wherein the plurality of fan blades is 8-20 fan blades.

[0146] Example 3. The turbomachinery engine according to any example herein, in particular example 1 or example 2, wherein the plurality of fan blades is 12-16 fan blades.

[0147] Example 4. The turbomachinery engine according to any example herein, in particular any of Examples 1-3, wherein the plurality of fan blades is exactly 12-14 fan blades.

[0148] Example 5. The turbomachinery engine according to any example herein, in particular any one of Examples 1-4, wherein RR is in the range of 0.125-0.55.

[0149] Example 6. The turbomachinery engine according to any example herein, in particular any one of Examples 1-5, wherein RR is in the range of 0.2-0.5.

[0150] Example 7. The turbomachinery engine according to any example herein, in particular any one of Examples 1-6, wherein RR is in the range of 0.25-0.35.

[0151] Example 8. The turbomachinery engine according to any example herein, in particular any one of Examples 1-6, wherein RR is in the range of 0.25-0.3.

[0152] Example 9. The turbomachinery engine according to any example herein, in particular any one of Examples 1-8, wherein the FPR at static sea level takeoff operating conditions is in the range of 1.05-1.5.

[0153] Example 10. The turbomachinery engine according to any example herein, in particular any one of Examples 1-9, wherein the FPR at static sea level takeoff operating conditions is in the range of 1.05-1.15.

[0154] Example 11. The turbomachinery engine according to any example herein, in particular any one of Examples 1-9, wherein the FPR at static sea level takeoff operating conditions is in the range of 1.2-1.4.

[0155] Example 12. The turbomachinery engine according to any example herein, in particular any of Examples 1-11, wherein F_Span at redline operating conditions is in the range of 20,000-200,000 lbf.

[0156] Example 13. The turbomachinery engine of any example herein, in particular any of Examples 1-12, wherein F_Span at redline operating conditions is in the range of 100,000-175,000 lbf.

[0157] Example 14. The turbomachinery engine according to any example herein, in particular any of Examples 1-12, wherein F_Span at redline operating conditions is in the range of 50,000-100,000 lbf.

[0158] Example 15. The turbomachinery engine according to any example herein, in particular Example 14, wherein F_Span at redline operating conditions is in the range of 65,000-85,000 lbf.

[0159] Example 16. A turbomachinery engine according to any example herein, in particular any one of Examples 1-15, wherein F_Area is between 3,000 and 25,000 in. 2 within the range.

[0160] Example 17. A turbomachinery engine according to any example herein, in particular any one of Examples 1-16, wherein F_Area is between 15,000 and 20,000 in. 2 within the range.

[0161] Example 18. A turbomachinery engine according to any example herein, in particular Example 17, wherein F_Area is between 17,000 and 18,000 in. 2 within the range.

[0162] Example 19. A turbomachinery engine according to any example herein, in particular any one of Examples 1-16, wherein F_Area is between 3,000 and 10,000 in. 2 within the range.

[0163] Example 20. A turbomachinery engine according to any example herein, in particular Example 18, wherein F_Area is between 4,000 and 7,000 in. 2 within the range.

[0164] Example 21. The turbomachinery engine according to any example herein, in particular any of Examples 1-20, wherein the fan blade tip speed of the fan assembly at redline operating conditions is in the range of 700-1,400 ft / s.

[0165] Example 22. The turbomachinery engine according to any example herein, and in particular Example 21, wherein the fan blade tip speed of the fan assembly at redline operating conditions is in the range of 800-950 ft / s.

[0166] Example 23. The turbomachinery engine according to any example herein, in particular Example 21, wherein the fan blade tip speed of the fan assembly at redline operating conditions is in the range of 1,000-1,200 ft / s.

[0167] Example 24. The turbomachinery engine according to any example herein, in particular any of Examples 1-23, wherein the gearbox comprises a gear ratio of 4.1-14, wherein the gear ratio is defined as the first rotational speed divided by the second rotational speed.

[0168] Example 25. A turbomachinery engine includes a fan assembly, a core engine, and a gearbox. The fan assembly includes a plurality of variable pitch fan blades, a first VPF parameter, and a second VPF parameter. The first VPF parameter is defined as the fan blade radius ratio (RR) divided by the fan pressure ratio (FPR) at static sea level takeoff operating conditions. The second VPF parameter is defined as the bearing span force (F_Span) measured in pounds-force divided by the fan area (F_Area) measured in square inches at redline operating conditions. The first VPF parameter is in the range of 0.1 to 0.25, and the second VPF parameter is in the range of 2-30 lbf / in 2 The core engine includes one or more compressor sections and one or more turbine sections. The gearbox includes an input and an output. The input is coupled to the one or more turbine sections of the core engine and has a first rotational speed, and the output is coupled to a fan assembly and has a second rotational speed less than the first rotational speed.

[0169] Example 26. The turbomachinery engine according to any example herein, in particular Example 25, further comprising a pitch change mechanism coupled to the plurality of variable pitch fan blades and configured to adjust the pitch of the plurality of variable pitch fan blades.

[0170] Example 27. The turbomachinery engine according to any example herein, in particular Example 26, wherein the pitch change mechanism is a linear drive pitch change mechanism.

[0171] Example 28. The turbomachinery engine according to any example herein, and in particular any of Examples 25-27, wherein the plurality of fan blades is 10-18 fan blades.

[0172] Example 29. The turbomachinery engine according to any example herein, in particular any of Examples 25-28, wherein the plurality of fan blades is 12-14 fan blades.

[0173] Example 30. The turbomachinery engine according to any example herein, in particular any of Examples 25-29, wherein the plurality of fan blades is exactly 14 fan blades.

[0174] Example 31. The turbomachinery engine according to any example herein, in particular any one of Examples 25-30, wherein RR is in the range of 0.125-0.55.

[0175] Example 32. The turbomachinery engine according to any example herein, in particular any one of Examples 25-31, wherein RR is in the range of 0.2-0.5.

[0176] Example 33. The turbomachinery engine according to any example herein, in particular any one of Examples 25-31, wherein RR is in the range of 0.25-0.35.

[0177] Example 34. The turbomachinery engine according to any example herein, in particular any one of Examples 25-33, wherein the FPR at static sea level takeoff operating conditions is in the range of 1.05-1.5.

[0178] Example 35. The turbomachinery engine according to any example herein, in particular any one of Examples 25-34, wherein the FPR at static sea level takeoff operating conditions is 1.05-1.15.

[0179] Example 36. The turbomachinery engine according to any example herein, in particular any of Examples 25-34, wherein the FPR at static sea level takeoff operating conditions is in the range of 1.20-1.40.

[0180] Example 37. The turbomachinery engine according to any example herein, in particular any of Examples 25-36, wherein F_Span at redline operating conditions is in the range of 20,000-200,000 lbf.

[0181] Example 38. The turbomachinery engine according to any example herein, in particular any of Examples 25-37, wherein F_Span at redline operating conditions is in the range of 60,000-90,000 lbf.

[0182] Example 39. The turbomachinery engine according to any example herein, in particular any of Examples 25-37, wherein F_Span at redline operating conditions is in the range of 100,000-150,000 lbf.

[0183] Example 40. A turbomachinery engine according to any example herein, in particular any one of Examples 25-39, wherein F_Area is between 3,000 and 25,000 in. 2 within the range.

[0184] Example 41. A turbomachinery engine according to any example herein, in particular any one of Examples 25-40, wherein F_Area is between 16,000 and 18,000 in. 2 within the range.

[0185] Example 42. A turbomachinery engine according to any example herein, in particular any one of Examples 25-40, wherein F_Area is between 5,500 and 6,500 in. 2 within the range.

[0186] Example 43. The turbomachinery engine according to any example herein, in particular any of Examples 25-42, wherein the fan blade tip speed of the fan assembly at redline operating conditions is in the range of 800-1,200 ft / s.

[0187] Example 44. The turbomachinery engine according to any example herein, in particular any of Examples 25-43, wherein the gearbox comprises a gear ratio of 6:1 to 11:1, and wherein the gear ratio is defined as the first rotational speed divided by the second rotational speed.

[0188] Example 45. The turbomachinery engine according to any example herein, in particular any one of Examples 25-44, wherein the second VPF parameter is in the range of 2.0-5.25 lbf / in 2 within the range.

[0189] Example 46. The turbomachinery engine according to any example herein, particularly any of Examples 25-44, wherein the second VPF parameter is between 5.25-30 lbf / in 2 within the range.

[0190] Example 47. A turbomachinery engine includes a fan assembly, a core engine, and a gearbox. The fan assembly includes a plurality of variable pitch fan blades, a first VPF parameter, and a second VPF parameter. The first VPF parameter is defined as the fan blade radius ratio (RR) divided by the fan pressure ratio (FPR) at static sea level takeoff operating conditions. The second VPF parameter is defined as the bearing span force (F_Span) measured in pounds-force divided by the fan area (F_Area) measured in square inches at redline operating conditions. The first VPF parameter is in the range of 0.10 to 0.40 and the second VPF parameter is in the range of 5.25-30 lbf / in 2 The core engine includes one or more compressor sections and one or more turbine sections. The gearbox includes an input and an output. The input is coupled to the one or more turbine sections of the core engine and has a first rotational speed, and the output is coupled to a fan assembly and has a second rotational speed less than the first rotational speed.

[0191] Example 48. The turbomachinery engine according to any example herein, in particular Example 47, further comprising a linearly driven pitch change mechanism coupled to the fan assembly and configured to adjust the pitch of the plurality of variable pitch fan blades.

[0192] Example 49. The turbomachinery engine according to any example herein, in particular any of Examples 47-48, wherein the plurality of fan blades is 12-16 fan blades.

[0193] Example 50. The turbomachinery engine according to any example herein, in particular any of Examples 47-49, wherein the plurality of fan blades is 12-14 fan blades.

[0194] Example 51. The turbomachinery engine according to any example herein, in particular any of Examples 47-50, wherein RR is in the range of 0.125-0.55.

[0195] Example 52. The turbomachinery engine according to any example herein, in particular any one of Examples 47-51, wherein RR is in the range of 0.2-0.5.

[0196] Example 53. The turbomachinery engine according to any example herein, in particular any one of Examples 47-51, wherein RR is in the range of 0.25-0.35.

[0197] Example 54. The turbomachinery engine according to any example herein, in particular any of Examples 47-53, wherein the FPR at static sea level takeoff operating conditions is in the range of 1.05-1.5.

[0198] Example 55. The turbomachinery engine according to any example herein, in particular any of Examples 47-54, wherein the FPR at static sea level takeoff operating conditions is in the range of 1.05-1.15.

[0199] Example 56. The turbomachinery engine according to any example herein, in particular any of Examples 47-54, wherein the FPR at static sea level takeoff operating conditions is in the range of 1.2-1.4.

[0200] Example 57. The turbomachinery engine according to any example herein, in particular any of Examples 47-56, wherein F_Span at redline operating conditions is in the range of 20,000-200,000 lbf.

[0201] Example 58. The turbomachinery engine according to any example herein, in particular any of Examples 47-57, wherein F_Span at redline operating conditions is in the range of 140,000-183,000 lbf.

[0202] Example 59. The turbomachinery engine according to any example herein, in particular any of Examples 47-57, wherein F_Span at redline operating conditions is in the range of 50,000-75,000 lbf.

[0203] Example 60. A turbomachinery engine according to any example herein, in particular any one of Examples 47-59, wherein F_Area is between 3,500 and 18,000 in. 2 within the range.

[0204] Example 61. A turbomachinery engine according to any example herein, in particular any one of Examples 47-60, wherein F_Area is between 15,000 and 18,000 in. 2 within the range.

[0205] Example 62. A turbomachinery engine according to any example herein, in particular any one of Examples 47-60, wherein F_Area is between 4,000 and 5,500 in. 2 within the range.

[0206] Example 63. A turbomachinery engine according to any example herein, in particular any of Examples 47-62, wherein the fan blade tip speed of the fan assembly under redline operating conditions is in the range of 800-950 ft / s.

[0207] Example 64. The turbomachinery engine according to any example herein, in particular any of Examples 47-63, wherein the gearbox comprises a gear ratio of 4.5:1 to 12:1.

[0208] Example 65. The turbomachinery engine according to any example herein, in particular any of Examples 47-64, wherein the first VPF parameter is in the range of 0.10 to 0.25.

[0209] Example 66. The turbomachinery engine according to any example herein, in particular any of Examples 47-64, wherein the first VPF parameter is in the range of 0.25 to 0.40.

[0210] Example 67. The turbomachinery engine according to any example herein, further comprising a third flow.

[0211] Example 68. A variable pitch fan assembly for a turbomachinery engine may be provided. The variable pitch fan assembly includes a plurality of variable pitch fan blades, a fan blade radius ratio (RR), a fan pressure ratio (FPR) at static sea level takeoff operating conditions, a bearing spanwise force measured in pounds-force at redline operating conditions (F_Span), and a fan area measured in square inches (F_Area). The variable pitch fan assembly is configured such that RR divided by FPR is in the range of 0.10-0.25, and F_Span divided by F_Area is in the range of 2.0-30.0 lbf / in 2 within the range.

[0212] Example 69. The variable pitch fan assembly according to any example herein, in particular Example 68, wherein F_Span divided by F_Area is in the range of 2-5.25 lbf / in. 2 within the range.

[0213] Example 70. The variable pitch fan assembly according to any example herein, in particular Example 68, wherein F_Span divided by F_Area is in the range of 5.25-30 lbf / in. 2 within the range.

[0214] Example 71. A variable pitch fan assembly for a turbomachinery engine may be provided. The variable pitch fan assembly includes a plurality of variable pitch fan blades, a fan blade radius ratio (RR), a fan pressure ratio (FPR) at static sea level takeoff operating conditions, a bearing spanwise force measured in pounds-force at redline operating conditions (F_Span), and a fan area measured in square inches (F_Area). The variable pitch fan assembly is configured such that RR divided by FPR is in the range of 0.10-0.40, and F_Span divided by F_Area is in the range of 5.25-30.0 lbf / in 2 within the range.

[0215] Example 72. A variable pitch fan assembly according to any example herein, in particular Example 71, wherein RR divided by FPR is in the range of 0.1-0.25.

[0216] Example 73. A variable pitch fan assembly according to any example herein, in particular Example 71, wherein RR divided by FPR is in the range of 0.25-0.40.

[0217] Example 74. A variable pitch fan assembly according to any example herein, wherein the fan blades are configured for use with an unducted engine.

[0218] Example 75. A variable pitch fan assembly according to any example herein, wherein the fan blades are configured for use with a ducted engine.

[0219] Example 76. The variable pitch fan assembly of any example herein, wherein the fan blades are configured for use with an engine including a third flow.

Claims

1. A turbomachinery engine, characterized in that: The turbomachinery engine comprises: A fan assembly, the fan assembly comprising a plurality of fan blades, a first VPF parameter and a second VPF parameter, wherein: The first VPF parameter is defined as the fan blade radius ratio (RR) divided by the fan pressure ratio (FPR) at static sea level takeoff operating conditions; The second VPF parameter is defined as the bearing span force (F_Span) measured in pounds-force divided by the fan area (F_Area) measured in square inches under redline operating conditions, where the bearing span force is defined as (mass of the fan blade / 386.4)*R_cg*ω 2 , where R_cg is the radius of the center of gravity of the fan blade measured in inches from the fan axis of rotation, and ω is the fan redline speed measured in radians per second; and The first VPF parameter is in the range of 0.1 to 0.25 and the second VPF parameter is in the range of 2-30 lbf / in 2 or the first VPF parameter is in the range of 0.1 to 0.4 and the second VPF parameter is in the range of 5.25-30 lbf / in 2 within the scope of a pitch changing mechanism coupled to the plurality of fan blades and configured to adjust a pitch of the plurality of fan blades; a blade assembly, the blade assembly comprising a plurality of blades disposed behind the fan blades; a core engine comprising one or more compressor sections and one or more turbine sections; and A gearbox includes an input and an output, wherein the input is coupled to the one or more turbine sections of the core engine and includes a first rotational speed, wherein the output is coupled to the fan assembly and has a second rotational speed less than the first rotational speed.

2. The turbomachinery engine according to claim 1, characterized in that: The plurality of fan blades is 8-20 fan blades.

3. The turbomachinery engine according to claim 1, characterized in that: The plurality of fan blades is 12-16 fan blades.

4. The turbomachinery engine according to claim 1, characterized in that: The plurality of fan blades is 12-14 fan blades.

5. The turbomachinery engine according to claim 1, characterized in that: The fan blade radius ratio is in the range of 0.125-0.

55.

6. The turbomachinery engine according to claim 1, characterized in that: The fan blade radius ratio is in the range of 0.2-0.

5.

7. The turbomachinery engine according to claim 1, characterized in that: The fan blade radius ratio is in the range of 0.25-0.

35.

8. A turbomachinery engine, characterized in that: The turbomachinery engine comprises: A fan assembly comprising a plurality of variable pitch fan blades, a first VPF parameter, and a second VPF parameter, wherein: The first VPF parameter is defined as the fan blade radius ratio (RR) divided by the fan pressure ratio (FPR) at static sea level takeoff operating conditions; The second VPF parameter is defined as the bearing span force (F_Span) measured in pounds-force divided by the fan area (F_Area) measured in square inches at redline operating conditions, where the bearing span force is defined as (mass of the variable pitch fan blade / 386.4)*R_cg*ω 2 , where R_cg is the radius of the center of gravity of the variable pitch fan blade measured in inches from the fan axis of rotation, and ω is the fan redline speed measured in radians per second; and The first VPF parameter is in the range of 0.1 to 0.25, and the second VPF parameter is in the range of 2-30 lbf / in 2 within the scope of a core engine comprising one or more compressor sections and one or more turbine sections; and A gearbox includes an input and an output, wherein the input is coupled to the one or more turbine sections of the core engine and includes a first rotational speed, wherein the output is coupled to the fan assembly and has a second rotational speed less than the first rotational speed.

9. The turbomachinery engine according to claim 8, characterized in that: Further included is a pitch change mechanism coupled to the plurality of variable pitch fan blades and configured to adjust the pitch of the plurality of variable pitch fan blades.

10. The turbomachinery engine according to claim 9, characterized in that: The pitch changing mechanism is a linear drive pitch changing mechanism.

11. The turbomachinery engine according to claim 8, characterized in that: Wherein the plurality of variable pitch fan blades is 10-18 fan blades.

12. The turbomachinery engine according to claim 8, characterized in that: Wherein the plurality of variable pitch fan blades is 12-14 fan blades.

13. The turbomachinery engine according to claim 8, characterized in that: Wherein the plurality of variable pitch fan blades is exactly 14 fan blades.

14. The turbomachinery engine according to claim 8, characterized in that: The fan blade radius ratio is in the range of 0.125-0.

55.

15. A turbomachinery engine, characterized in that: The turbomachinery engine comprises: A fan assembly comprising a plurality of variable pitch fan blades, a first VPF parameter, and a second VPF parameter, wherein: The first VPF parameter is defined as the fan blade radius ratio (RR) divided by the fan pressure ratio (FPR) at static sea level takeoff operating conditions; The second VPF parameter is defined as the bearing span force (F_Span) measured in pounds-force divided by the fan area (F_Area) measured in square inches at redline operating conditions, where the bearing span force is defined as (mass of the variable pitch fan blade / 386.4)*R_cg*ω 2 , where R_cg is the radius of the center of gravity of the variable pitch fan blade measured in inches from the fan axis of rotation, and ω is the fan redline speed measured in radians per second; and The first VPF parameter is in the range of 0.10 to 0.40, and the second VPF parameter is in the range of 5.25-30.0 lbf / in 2 within the scope of a core engine comprising one or more compressor sections and one or more turbine sections; and A gearbox includes an input and an output, wherein the input is coupled to the one or more turbine sections of the core engine and includes a first rotational speed, wherein the output is coupled to the fan assembly and has a second rotational speed less than the first rotational speed.

16. The turbomachinery engine according to claim 15, characterized in that Further included is a linear-drive pitch change mechanism coupled to the fan assembly and configured to adjust the pitch of the plurality of variable-pitch fan blades.

17. The turbomachinery engine according to claim 15, characterized in that: Wherein the plurality of variable pitch fan blades is 12-16 fan blades.

18. The turbomachinery engine according to claim 15, characterized in that: Wherein the plurality of variable pitch fan blades is 12-14 fan blades.

19. The turbomachinery engine according to claim 15, characterized in that: The fan blade radius ratio is in the range of 0.125-0.

55.

20. The turbomachinery engine according to claim 15, characterized in that The fan blade radius ratio is in the range of 0.2-0.5.

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