阅读说明：本技术 涡轮机以及具有轴向偏移太阳齿轮和环形齿轮的周转齿轮组件 (Turbine and epicyclic gear assembly with axially offset sun and ring gears ) 是由 马可·法基尼 詹卢卡·安德烈 保罗·阿尔塔穆拉 于 2021-05-06 设计创作，主要内容包括：提供了一种涡轮机发动机,该涡轮机发动机包括风扇组件和核心发动机,该核心发动机包括涡轮和可与涡轮一起旋转的输入轴。单级周转齿轮组件以第一速度接收输入轴,并以第二速度驱动联接至风扇组件的输出轴。太阳齿轮绕齿轮组件的纵向中心线旋转,并且具有沿着齿轮组件的纵向中心线的太阳齿轮啮合区域,太阳齿轮被构造成在太阳齿轮啮合区域接触多个行星齿轮。环形齿轮啮合区域沿着齿轮组件的纵向中心线设置,环形齿轮被构造成在环形齿轮啮合区域接触多个行星齿轮。太阳齿轮啮合区域沿着纵向中心线与环形齿轮啮合区域轴向偏移。(A turbine engine is provided that includes a fan assembly and a core engine including a turbine and an input shaft rotatable with the turbine. A single-stage epicyclic gear assembly receives the input shaft at a first speed and drives an output shaft coupled to the fan assembly at a second speed. The sun gear rotates about a longitudinal centerline of the gear assembly and has a sun gear mesh region along the longitudinal centerline of the gear assembly where the sun gear is configured to contact the plurality of planet gears. A ring gear mesh region is disposed along a longitudinal centerline of the gear assembly, the ring gear being configured to contact the plurality of planet gears at the ring gear mesh region. The sun gear mesh region is axially offset from the ring gear mesh region along the longitudinal centerline.)

1. A turbomachine engine (100, 600), comprising:

a fan assembly (104), the fan assembly (104) including a plurality of fan blades (108); and

a core engine (106), the core engine (106) including a turbine and an input shaft (146) rotatable with the turbine; and

a single-stage epicyclic gear assembly (102, 202, 302, 402, 502), the single-stage epicyclic gear assembly (102, 202, 302, 402, 502) receiving the input shaft (146) at a first speed and driving an output shaft (124) coupled to the fan assembly (104) at a second speed, the second speed being slower than the first speed, the gear assembly (102, 202, 302, 402, 502) comprising:

a sun gear (204, 304, 404, 504), a plurality of planet gears (206, 306, 406, 506) and a ring gear (208, 308, 408, 508), the sun gear rotating about a longitudinal centerline (120) of the gear assembly;

a sun gear engagement region (228, 230, 328, 330, 428, 430), the sun gear engagement region (228, 230, 328, 330, 428, 430) being along the longitudinal centerline (120) of the gear assembly, the sun gear (204, 304, 404, 504) being configured to contact the plurality of planet gears (206, 306, 406, 506) at the sun gear engagement region (228, 230, 328, 330, 428, 430);

a ring gear engagement region (232, 234, 332, 334, 432, 434), the ring gear engagement region (232, 234, 332, 334, 432, 434) being along the longitudinal centerline (120) of the gear assembly, the ring gear (208, 308, 408, 508) being configured to contact the plurality of planet gears (206, 306, 406, 506) at the ring gear engagement region (232, 234, 332, 334, 432, 434),

wherein the sun gear mesh region (228, 230, 328, 330, 428, 430) is axially offset from the ring gear mesh region (232, 234, 332, 334, 432, 434) along the longitudinal centerline (120) such that at least 50% of the width of the sun gear mesh region does not axially overlap the ring gear mesh region.

2. The turbomachine engine of claim 1, wherein the sun gear (204, 304, 404, 504), the plurality of planet gears (206, 306, 406, 506) and the ring gear (208, 308, 408, 508) comprise double bevel gears, and the sun gear (204, 304, 404, 504) comprises a first sun gear set (210, 310, 410, 510) and a second sun gear set (212, 312, 412, 512), each of the plurality of planet gears (206, 306, 406, 506) comprises a first planetary gear set (214, 314, 414, 514) and a second planetary gear set (216, 316, 418, 518), and the ring gear (208, 308, 408, 508) comprises a first ring gear set (218, 318, 418, 518) and a second ring gear set (220, 320, 420, 520).

3. The turbomachine engine of claim 2, wherein the first and second ring gear sets (218, 318, 418, 518, 220, 320, 420, 520) are axially spaced from one another along the longitudinal centerline, and the first and second sun gear sets (210, 310, 410, 510, 212, 312, 412, 512) are positioned between the first and second ring gear sets (218, 318, 418, 518, 220, 320, 420, 520).

4. The turbomachine engine according to claim 2, wherein the first sun gear set (210, 310, 410, 510) and the second sun gear set (212, 312, 412, 512) are axially spaced from each other along the longitudinal centerline (120), and the first ring gear set (218, 318, 418, 518) and the second ring gear set (220, 320, 420, 520) are positioned between the first sun gear set (210, 310, 410, 510) and the second sun gear set (212, 312, 412, 512).

5. The turbomachine engine according to any one of claims 2 to 4, wherein the sun gear mesh region comprises a first sun gear mesh region (228, 328, 428) and a second sun gear mesh region, the first sun gear set (210, 310, 410, 510) meshing with the first planetary gear set (214, 314, 414, 514) at the first sun gear mesh region (228, 328, 428), the second sun gear set (212, 312, 412, 512) meshing with the second planetary gear set (216, 316, 418, 518) at the second sun gear mesh region, and

the ring gear engagement zones include a first ring gear engagement zone where the first ring gear set (218, 318, 418, 518) is in mesh with the first planetary gear set (214, 314, 414, 514), and a second ring gear engagement zone where the second ring gear set (220, 320, 420, 520) is in mesh with the second planetary gear set (216, 316, 418, 518).

6. The turbomachine engine of claim 5, wherein the first sun gear mesh region and the first ring gear mesh region do not axially overlap along the longitudinal centerline.

7. The turbomachine engine of claim 5, wherein an axial gap exists between the first sun gear mesh region and the first ring gear mesh region.

8. The turbomachine engine of claim 7, wherein a gap width of the axial gap is less than 15% of a width of the first planetary gear set (214, 314, 414, 514), less than 10% of the width of the first planetary gear set, less than 5% of the width of the first planetary gear set, or less than 2% of the width of the first planetary gear set.

9. The turbomachine engine of claim 5, wherein there is an axial overlap between the first sun gear mesh region and the first ring gear mesh region, and an amount of the axial overlap is less than 15% of a width of the first planetary gear set, less than 10% of the width of the first planetary gear set, less than 5% of the width of the first planetary gear set, or less than 2% of the width of the first planetary gear set.

10. The turbomachine of any one of the preceding claims, wherein the gear ratio of the gear assembly ranges from 6: 1 to 14: 1, from 6.6 to 12: 1, from 7: 1 to 12: 1 or from 8: 1 to 12: 1.

Technical Field

The present subject matter relates generally to turbomachines including gear assemblies, and more particularly to gear assembly arrangements specific to certain turbomachinery configurations.

Thank you government support

The project that led to the application had received funds from clear Sky 2Joint Underwarking (JU) according to the withdrawal agreement No 945541. JU was supported by european union horizon 2020 research and innovation programs and Clean Sky 2JU members other than the european union.

Background

Turbofan engines operate on the principle that a central gas turbine core drives a bypass fan located at a radial position between the nacelle and the engine core. With this configuration, the engine is typically limited by the allowable size of the bypass fan, since increasing the size of the fan correspondingly increases the size and weight of the nacelle.

In contrast, open rotor engines operate on the principle of having the bypass fan located outside the engine compartment. This allows the use of larger rotor blades that can act on a larger air volume than conventional turbofan engines, potentially improving propulsion efficiency over conventional turbofan engine designs.

Turbine engine designs, including turbofan and open rotor engines, may require a large gear ratio between the low speed spool and the fan rotor to allow larger rotor blades to act on a larger air volume and/or do so at certain desired operating speeds of the engine or aircraft. One challenge is that known gear assemblies may not provide sufficient gear ratios for desired operation. For example, known gear assemblies may not adequately reduce the output rotational speed relative to the input rotational speed such that the fan rotor operates too fast and inefficiently and/or the turbine operates too slow and inefficiently.

As such, there is a need to provide a gear assembly that may be adapted to the desired gear ratio of certain turbine configurations.

Disclosure of Invention

Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the technology disclosed in the description.

Various turbine engines and gear assemblies are disclosed herein. In some embodiments, a turbine engine is provided that includes a fan assembly and a core engine including a turbine and an input shaft rotatable with the turbine. The single-stage epicyclic gear assembly receives the input shaft at a first speed and drives an output shaft coupled to the fan assembly at a second speed, the second speed being slower than the first speed. The gear assembly includes a sun gear, a plurality of planet gears, and a ring gear. The sun gear rotates about a longitudinal centerline of the gear assembly and has a sun gear mesh region along the longitudinal centerline of the gear assembly where the sun gear is configured to contact the plurality of planet gears. A ring gear mesh region is disposed along a longitudinal centerline of the gear assembly, the ring gear being configured to contact the plurality of planet gears at the ring gear mesh region. The sun gear mesh region is axially offset from the ring gear mesh region along an axial centerline.

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 embodiments of the disclosed technology and together with the description, serve to explain the principles of the disclosure.

Drawings

A full and enabling disclosure of the present invention, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which:

FIG. 1 is a cross-sectional schematic view of an exemplary embodiment of an open rotor propulsion system;

FIG. 2 is a cross-sectional schematic view of an exemplary embodiment of an open rotor propulsion system;

FIG. 3 is a diagrammatical representation of an alternative embodiment of an exemplary vane assembly for an open rotor propulsion system;

FIG. 4 is a schematic view of an exemplary gear assembly having axially offset gear mesh regions;

FIG. 5 is a schematic view of an exemplary gear assembly having axially offset gear mesh regions;

FIGS. 6A and 6B are schematic views of a gear assembly having three planetary gears;

FIGS. 7A and 7B are schematic views of a gear assembly having two planetary gears;

FIG. 8 is a schematic view of an exemplary gear assembly having axially offset gear mesh regions;

FIG. 9 is a schematic view of an exemplary gear assembly having axially offset gear mesh regions;

FIG. 10 is a schematic view of an exemplary gear assembly having axially offset gear mesh regions with no overlap;

FIG. 11 is a schematic view of an exemplary gear assembly having axially offset gear mesh regions with overlap; and

FIG. 12 is a cross-sectional schematic view of an exemplary embodiment of a ducted propulsion system.

Detailed Description

Reference now will be made in detail to embodiments of the invention, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation of the invention, not limitation of the invention. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present invention without departing from the scope or spirit of the invention. For instance, features illustrated or described as part of one embodiment, can be used with another embodiment to yield a still further embodiment. Thus, it is intended that the present invention cover the modifications and variations of this invention provided they come within the scope of the appended claims and their equivalents.

The word "exemplary" is used herein to mean "serving as an example, instance, or illustration. Any embodiment described herein as "exemplary" is not necessarily to be construed as preferred or advantageous over other embodiments.

As used herein, the terms "first," "second," and "third" may be used interchangeably to distinguish one element from another, and are not intended to denote the position or importance of the various elements.

The terms "forward" and "aft" refer to relative positions within the gas turbine engine or vehicle and to normal operating attitudes of the gas turbine engine or vehicle. For example, for a gas turbine engine, forward refers to a position closer to the engine inlet, and aft refers to a position closer to the engine nozzle or exhaust outlet.

The terms "upstream" and "downstream" refer to relative directions with respect to fluid flow in a fluid path. For example, "upstream" refers to the direction from which the fluid flows, while "downstream" refers to the direction to which the fluid flows.

Unless specified otherwise, the terms "coupled," "secured," "attached," and the like refer to a direct coupling, securing, or attachment, as well as an indirect coupling, securing, or attachment through one or more intermediate components or features.

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

Approximating language, as used herein throughout the specification and claims, is applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms, such as "about", "about" and "substantially", are not to be limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of an instrument for measuring the value, or the precision of a method or machine for constructing or manufacturing the component and/or system. For example, approximating language may refer to within a margin of 1%, 2%, 4%, 10%, 15%, or 20% of an individual value, range of values, and/or endpoint of a defined range of values.

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

One or more components of the turbine engine or gear assembly described below may be manufactured or formed using any suitable process, such as an additive manufacturing process, such as a 3D printing process. Such components may be integrally formed as a single integral component, or any suitable number of sub-components, using such processes. In particular, additive manufacturing processes may allow such components to be integrally formed and include various features that are not possible when using existing manufacturing methods. For example, the additive manufacturing methods described herein enable the manufacture of heat exchangers having unique features, configurations, thicknesses, materials, densities, fluid passages, headers, and mounting structures that may not be possible or practical using existing manufacturing methods. Some of these features are described herein.

Referring now to the drawings, fig. 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, core engine 106 is a Brayton cycle system configured to drive fan assembly 104. Core engine 106 is at least partially enclosed by a casing 114. Fan assembly 104 includes a plurality of fan blades 108. The bucket assembly 110 extends from the casing 114. A vane assembly 110 including a plurality of vanes 112 is positioned in operable arrangement with fan blades 108 to provide thrust relative to fan blades 108, control thrust vectors, attenuate or redirect undesirable acoustic noise, and/or otherwise desirably alter airflow. In some embodiments, the fan assembly 104 includes between three (3) and twenty (20) fan blades 108. In a particular embodiment, the fan assembly 104 includes between ten (10) and sixteen (16) fan blades 108. In certain embodiments, fan assembly 104 includes twelve (12) fan blades 108. In certain embodiments, the bucket assembly 110 includes an equal or lesser number of buckets 112 as fan blades 108.

In some embodiments, fan blade tip speeds may reach 650 to 900fps, or 700 to 800fps at cruise flight conditions. The Fan Pressure Ratio (FPR) of the fan assembly 104 may be 1.04 to 1.10, or 1.05 to 1.08 in some embodiments, as measured across the fan blades at cruise flight conditions. In some embodiments, the output torque provided by the gear assembly may be in the range of 20kNm to 200kNm, or in other embodiments in the range of 40kNm to 150 kNm.

In certain embodiments, such as shown in FIG. 1, the vane assembly 110 is positioned downstream or aft of the fan assembly 104. However, it should be understood that in some embodiments, the bucket assembly 110 may be positioned upstream or forward of the fan assembly 104. In still various embodiments, the engine 100 may include a first vane assembly positioned forward of the fan assembly 104 and a second vane assembly positioned aft of the fan assembly 104. The fan assembly 104 may be configured to desirably adjust the pitch (pitch) at one or more fan blades 108, for example, to control the thrust vector, attenuate or redirect noise, and/or vary the thrust output. The bucket assemblies 110 may be configured to desirably adjust the pitch at one or more of the buckets 112, for example, to control thrust vectors, attenuate or redirect noise, and/or vary thrust output. The pitch control mechanisms at one or both of the fan assembly 104 or the vane assembly 110 may cooperate to produce one or more of the desired effects described above.

In certain embodiments, such as shown in FIG. 1, engine 100 is a non-ducted thrust producing system such that the plurality of fan blades 108 are not shrouded by the nacelle or fan casing. As such, in various embodiments, the engine 100 may be configured as a non-shrouded turbofan engine, an open rotor engine, or a paddle fan engine. In a particular embodiment, engine 100 is a single non-ducted rotary engine that includes a single row of fan blades 108. The engine 100 configured as an open rotor engine includes a fan assembly 104 having large diameter fan blades 108, such as may be suitable for high bypass ratios, high cruise speeds (e.g., comparable to, or generally higher cruise speeds than, an aircraft having turbofan engines), high cruise altitudes (e.g., comparable to, or generally higher cruise speeds than, an aircraft having turbofan engines), and/or relatively low rotational speeds. Cruise altitude generally refers to the level at which an aircraft is after climbing and before descending to an approach flight phase. In various embodiments, the engine is applied to a vehicle that is cruising at a height of up to about 65,000 ft. In certain embodiments, the cruising height is between about 28,000ft to about 45,000 ft. Still in certain embodiments, cruise altitude is expressed in Flight Level (FL) based on standard barometric pressure at sea level, where cruise flight conditions are between FL280 and FL 650. In another embodiment, the cruise flight conditions are between FL280 and FL 450. In still other embodiments, the cruising altitude is defined based at least on atmospheric pressure, wherein the cruising altitude is between about 4.85psia and about 0.82psia based on a sea level pressure of about 14.70psia and a sea level temperature of about 59 degrees Fahrenheit. In another embodiment, the cruising height is between about 4.85psia and about 2.14 psia. It should be appreciated that in some embodiments, the range of cruising heights defined by pressures may be adjusted based on different reference sea level pressures and/or sea level temperatures.

Core engine 106 is typically enclosed in a casing 114 that defines a maximum diameter. In certain embodiments, the engine 100 includes a length from a longitudinal forward end 116 to a longitudinal rearward end 118. In various embodiments, the engine 100 defines a length (L) and a maximum diameter (D)_max) To provide reduced installation resistance. In one embodiment, the L/D_maxIs at least 2. In another embodiment, L/D_maxIs at least 2.5. In some embodiments, the L/D_maxLess than 5, less than 4 and less than 3. In various embodiments, it should be understood that L/D_maxFor a single non-ducted rotary engine.

The reduced installation resistance may further provide improved efficiency, such as improved specific fuel consumption. Additionally or alternatively, the reduced drag may provide cruise altitude engine and aircraft operation at or above 0.5 mach. In certain embodiments, the L/D_maxThe fan assembly 104 and/or the vane assembly 110, respectively or together, at least partially configure the engine 100 to operate at a maximum cruise altitude operating speed between about Mach 0.55 and about Mach 0.85.

Referring again to fig. 1, the core engine 106 extends in a radial direction R relative to an engine axis centerline 120. The gear assembly 102 receives power or torque from the core engine 106 through a power input source 122 and provides power or torque to drive the fan assembly 104 in a circumferential direction C about the engine axis centerline 120 through a power output source 124.

FIG. 2 illustrates a front cross-sectional view of an exemplary embodiment of open rotor propulsion engine 100. The engine 100 has a fan assembly 104, the fan assembly 104 including a plurality of fan blades 108 about a central longitudinal axis 120 of the engine 100. The fan blades 108 are circumferentially arranged at equal intervals about the centerline 120, and each fan blade 108 has a root 125, a tip 126, an axial span defined therebetween, and a center blade axis 128.

Core engine 16 includes a compressor section 130, a heat addition system 132 (e.g., a combustor), and an expansion section 134 arranged together in a serial flow arrangement. The core engine 106 extends circumferentially relative to an engine centerline axis 120. Core engine 106 includes a high speed spool that includes a high speed compressor 136 and a high speed turbine 138 operatively rotatably coupled together by a high speed shaft 140. The heat addition system 132 is positioned between a high speed compressor 136 and a high speed turbine 138. Various embodiments of the heat addition system 132 include a combustion section. The combustion section may be configured as a deflagration combustion section, a rotary detonation combustion section, a pulse detonation combustion section, or other suitable heat addition system. The heat addition system 132 may be configured as one or more of a rich or lean burn system, or a combination thereof. In still various embodiments, the heat addition system 132 includes an annular combustor, a can combustor, a sleeve combustor, a Trapped Vortex Combustor (TVC), or other suitable combustion system or combination thereof.

Core engine 106 also includes a booster or low speed compressor positioned in flow relationship with high speed compressor 136. The low-speed compressor 142 is rotatably coupled to a low-speed turbine 144 via a low-speed shaft 146 such that the low-speed turbine 144 is capable of driving the low-speed compressor 142. Low-speed shaft 146 is also operably connected to gear assembly 102 to provide power to fan assembly 104, as further described herein.

It should be understood that, unless otherwise specified, the terms "low" and "high," or their respective comparative stages (e.g., lower, higher, if applicable) when used with a compressor, turbine, shaft or spool component, refer to the relative speeds within the engine. For example, "low turbine" or "low speed turbine" defines a component configured to operate at a lower rotational speed (e.g., maximum allowable rotational speed) than a "high turbine" or "high speed turbine" of the engine. Alternatively, the foregoing terms may be understood at their highest level unless otherwise specified. For example, "low turbine" or "low speed turbine" may refer to the lowest maximum rotational speed turbine within the turbine section, "low compressor" or "low speed compressor" may refer to the lowest maximum rotational speed turbine within the compressor section, "high turbine" or "high speed turbine" may refer to the highest maximum rotational speed turbine within the turbine section, and "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 the terms "low" or "high" in the above-described aspects may additionally or alternatively be understood as a minimum or maximum allowable speed relative to a minimum allowable speed, or relative to normal, desired, steady state, etc. operation of the engine.

As discussed in more detail below, the core engine 106 includes a gear assembly 102, the gear assembly 102 configured to transfer power from the extension section 140 and reduce an output rotational speed at the fan assembly 104 relative to the low speed turbine 144. The embodiments of the gear assembly 104 depicted and described herein may allow for a gear ratio suitable for large diameter non-ducted fans. Additionally, embodiments of the gear assembly 102 provided herein may be suitable within the radial or diametric constraints of the core engine 106 within the casing 114.

In the exemplary embodiment of FIG. 2, engine 100 also includes a bucket assembly 110, where bucket assembly 110 includes a plurality of buckets 112 disposed about a central axis 120. Each bucket 112 has a root 148 and a tip 150, and a span defined therebetween. The vanes 112 may be arranged in various ways. For example, in some embodiments, they are not all equidistant from the rotating assembly.

In some embodiments, the vanes 112 are mounted to a fixed frame and do not rotate relative to the central axis 120, but may include mechanisms for adjusting their orientation relative to their axes 154 and/or relative to the blades 108. For reference purposes, fig. 2 depicts a forward direction, indicated by arrow F, which in turn defines forward and rearward portions of the system. As shown in fig. 1 and 2, the fan assembly 104 may be located in a "pull" configuration forward of the gas core engine 106, while the exhaust port 156 is located aft of the core engine 106.

Left-or right-turn engine configurations, which may be useful for certain facilities in reducing the effects of multi-engine torque on an aircraft, may be achieved by mirroring airfoils (e.g., 108, 112) such that the fan assembly 104 rotates clockwise for one propulsion system and counterclockwise for the other propulsion system. Alternatively, an optional reversing gearbox may be provided to allow the fan blades to be rotated clockwise or counterclockwise, i.e., to provide a left-or right-hand configuration as desired, using a common gas turbine core and low pressure turbine, e.g., a pair of counter-rotating engine assemblies may be provided for certain aircraft facilities without having internal engine parts designed for counter-rotating directions.

The engine 100 also includes a gear assembly 102, the gear assembly 102 including a gear set for reducing the rotational speed of the fan assembly 104 relative to the low speed (pressure) turbine 144. In operation, the rotating fan blades 108 are driven by the low speed (pressure) turbine 144 via the gear assembly 102 such that the fan blades 108 rotate about the axis 120 and generate thrust to propel the engine 100, thereby propelling the aircraft on which the engine 100 is mounted, in the forward direction F.

It may be desirable for one or both of the fan blades 104 or the vanes 112 to include a pitch mechanism such that the blades may rotate independently or in conjunction with each other relative to a pitch axis of rotation (labeled 128 or 154, respectively). Such pitch changes may be used to alter thrust and/or vortex effects under various operating conditions, including providing a thrust reversal feature, which may be useful under certain operating conditions (e.g., while the aircraft is landing).

The vanes 112 may be sized and configured to impart a canceling (counter) swirl to the fluid such that the swirl of the fluid in the downstream direction behind the fan blades 104 and vanes 112 is substantially reduced, which translates into an increased level of induced efficiency. As shown in fig. 1 and 2, the vanes 112 may have a shorter span than the fan blades 104. For example, the span of the vanes 112 is at least 50% of the span of the fan blades 104. In some embodiments, the span of the vanes may be the same or longer than the span of the fan blades 104, if desired. As shown in fig. 1, vanes 112 may be attached to an aircraft structure associated with engine 100, or another aircraft structure such as a wing, a pylon, or a fuselage. The number of vanes 112 may be less than, greater than, or equal to the number of fan blades 104. In some embodiments, the number of vanes 112 is greater than two or greater than four. The fan blade 104 may be sized, shaped, and contoured to account for the desired blade loading.

In the embodiment shown in FIG. 2, an annular 360 degree inlet 158 is located between the fan assembly 104 and the vane assembly 110 and provides a path for incoming atmospheric air to enter the engine core 106 radially inward from at least a portion of the vane assembly 110. Such a location may be advantageous for various reasons, including managing icing performance and protecting inlet 158 from various objects and materials that may be encountered in operation.

Fig. 1 and 2 illustrate a so-called "pull" configuration, in which the fan assembly 104 is located forward of the engine core 106. Other configurations are also possible and contemplated within the scope of the present disclosure, such as so-called "push" configuration embodiments in which the engine core 106 is located forward of the fan assembly 104.

The selection of a "pull" or "push" configuration may be made consistent with the selection of an installation orientation relative to the fuselage of the intended aircraft application, and some choices may be structurally or operationally advantageous depending on whether the installation location and orientation is wing-mounted, fuselage-mounted, or tail-mounted configuration.

In the exemplary embodiment of FIG. 2, in addition to an open rotor or non-ducted fan assembly 104 having a plurality of fan blades 104, an optional ducted fan 160 is included behind the fan assembly 104, such that the engine 100 includes both ducted and non-ducted fans, both for generating thrust by air movement at atmospheric temperature without passing through the engine core 106. Ducted fan 160 is shown at about the same axial location as vanes 112, and radially inward of vane root 148. Alternatively, ducted fan 160 may be between vanes 112 and core duct 162, or further forward of vanes 112. The ducted fan 160 may be driven by a low pressure turbine or by any other suitable source of rotation, and may serve as the first stage of the booster 142 or may operate separately. Air entering the inlet 158 flows through the inlet duct 164 and is then split such that a portion flows through the core duct 162 and a portion flows through the fan duct 166. Fan duct 166 may incorporate a heat exchanger 168 and exhaust to atmosphere through a separate fixed or variable nozzle 170 aft of bucket assembly 110, nozzle 170 being outboard of fan casing 152 aft end and engine core casing 172. Thus, air flowing through the fan duct 166 "bypasses" the core of the engine, rather than passing through the core.

Thus, in the exemplary embodiment, engine 100 includes a non-ducted fan formed by fan blades 108 followed by a ducted fan 160 that channels airflow into two concentric or non-concentric ducts 162 and 166, forming a three-flow engine architecture having 3 paths for air passing through fan assembly 104.

In the exemplary embodiment shown in fig. 2, a slidable, movable and/or translatable plug nozzle 172 having an actuator may be included to vary the exit area of the nozzle 170. Plug nozzles are generally annular symmetrical devices that adjust the opening area of an outlet (e.g., fan flow or core flow) by axial movement of the nozzle such that the gap between the nozzle surface and a stationary structure (e.g., adjacent walls of a duct) varies in a predetermined manner to reduce or increase the space of airflow through the duct. Other suitable nozzle designs may also be employed, including those incorporating a thrust reversal function. 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 to have a failure mode (e.g., fully open, fully closed, or intermediate positions) such that the nozzle 170 has a consistent "home" position that returns in the event of any system failure, which may prevent commands from reaching the nozzle 170 and/or its actuators.

In some embodiments, a mixing device 174 may be included in the area behind the core nozzle 176 to help mix the fan flow and the core flow to improve acoustic performance by directing the core flow outward and the fan flow inward.

Because the engine 100 shown in FIG. 2 includes the open rotor fan assembly 104 and the ducted fan assembly 160, the thrust outputs of the two and the work distribution therebetween may be tailored to achieve specific thrust, fuel combustion, thermal management, and/or acoustic characterization goals that may be superior to those of a typical ducted fan gas turbine propulsion assembly of comparable thrust class. By reducing the proportion of thrust required to be provided by the non-ducted fan assembly 104, the ducted fan assembly 160 may allow for a reduction in the overall fan diameter of the non-ducted fan assembly, thereby providing installation flexibility and reduced weight.

In operation, the engine 100 may include a control system that manages the loading of the respective open and ducted fans, and potentially the exit area of the variable fan nozzle, to provide different thrust, noise, cooling capability and other performance characteristics for various portions of the flight envelope and various operating conditions associated with aircraft operation. For example, in climb mode, the ducted fan may be operated at a maximum pressure ratio to maximize the thrust capacity of the flow, while in cruise mode, the ducted fan may be operated at a lower pressure ratio to increase overall efficiency by relying on thrust from non-ducted fans. Nozzle actuation adjusts the ducted fan operating line and the total engine fan pressure ratio regardless of total engine airflow.

The ducted fan flow through the fan duct 166 may include one or more heat exchangers 168 for removing heat from various fluids used in engine operation, such as air-cooled oil coolers (ACOC), Cooled Cooling Air (CCA), and the like. The heat exchanger 168 may take advantage of integration with the fan duct 166, compared to conventional ducted fan architectures, with reduced performance losses (e.g., fuel efficiency and thrust) because the primary source of thrust, in this case non-ducted fan flow, is not affected. The heat exchanger may cool a fluid, such as gearbox oil, engine sump oil, a heat transfer fluid (e.g., a supercritical fluid) or a commercially available single or two-phase fluid (supercritical CO2, EGV, Slither 800, liquid metal, etc.), engine bleed air, etc. The heat exchanger may also consist of different sections or channels (e.g., ACOC paired with a fuel cooler) that cool different working fluids. The heat exchanger 168 may 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 the heat to the heat exchanger.

Since the fan pressure ratio of ducted fans is higher than the fan pressure ratio of non-ducted fans, the fan duct provides an environment in which a more compact heat exchanger can be utilized than outside of the core cowl installed in the non-ducted fan flow. The Fan Pressure Ratio (FPR) of the fan bypass air is very low (1.05 to 1.08) and it is therefore difficult to drive the air through the heat exchanger. Without a fan duct as described herein, a scoop or booster bleed air may be required to provide cooling air to and through the heat exchanger. A set of parameters may be established around the heat exchanger in the fan duct based on the thermal load, heat exchanger size, ducted fan flow correction flow, and ducted fan flow temperature.

The fan duct 166 also provides other advantages in terms of reduced nacelle drag, enabling more aggressive nacelle closure, improved core flow particle separation, and inclement weather operation. By discharging the fan duct flow over the core cowl, this helps to energize the boundary layer and enables selection of a steeper nacelle closing angle between the maximum size of the core cowl 172 and the exhaust port 156. The closing angle is generally limited by airflow separation, but boundary layer excitation by air discharged on the core cowl from fan duct 166 reduces airflow separation. This results in a shorter, lighter structure with less frictional surface resistance.

The fan assembly 104 and/or the vane assembly may be shrouded or unshrouded (as shown in fig. 1 and 2). FIG. 3 shows an optional annular shroud or duct 178 on the bucket assembly 110 distal to the axis 120. In addition to the benefits of noise reduction, the conduit 178 shown in FIG. 3 may also provide improved vibrational response and structural integrity of the stationary vanes 112 by coupling the stationary vanes 112 into an assembly that forms an annular ring or one or more circumferential sectors (sectors), i.e., segments that form portions of an annular ring linking two or more vanes 112. The conduit 178 may also allow the pitch of the vanes to be more easily changed. FIG. 12, discussed in more detail below, discloses another embodiment in which both the fan assembly and the vane assembly are shrouded.

Embodiments of gear assembly 102 depicted and described herein may provide an L/D suitable for engine 10_maxGear ratios and arrangements within the constraints. In certain embodiments, the gear assembly depicted and described with respect to fig. 4-11 allows for providing gear ratios and arrangements of rotational speeds of the fan assembly corresponding to one or more ranges of cruising altitude and/or cruising speed provided above.

Various embodiments of the gear assembly 102 provided herein may allow for up to 14: a gear ratio of 1. Still other various embodiments of the gear assembly 102 provided herein can allow for a gear ratio of at least 6: 1. Still further various embodiments of the gear assembly 102 provided herein allow for a gear ratio between 6: 1 and 12: 1 for a single-stage epicyclic gear assembly. It should be appreciated that embodiments of the gear assembly 102 provided herein may allow for large gear ratios and within constraints (e.g., without limitation, length (L) of engine 10, maximum diameter (D) of engine 100)_max) A cruising altitude of up to 65,000ft, and/or an operating cruising speed of up to mach 0.85, or combinations thereof).

Various example gear assemblies are shown and described herein. These gear assemblies may be used with any of the exemplary engines and/or any other suitable engine that may require such gear assemblies. In this manner, it will be appreciated that the gear assemblies disclosed herein are generally operable with an engine having a rotating element with a plurality of rotor blades and a turbine having a turbine and a shaft rotatable with the turbine. For such engines, the rotating element (e.g., fan assembly 104) may be driven by a shaft of the turbine (e.g., low speed shaft 146) through a gear assembly.

Fig. 4 illustrates an example gear assembly 202 having an axially offset face width. The gear assembly 202 includes a diameter D_sSun gear 204 of diameter D_pA plurality of planet gears 206 and a diameter D_rRing gear 208. Each of the sun gear 204, the planet gears 206, and the ring gear 208 is a double helical gear (double helical gear) having first and second sets of helical teeth that are inclined at an acute angle relative to each other. In particular, the sun gear 204 includes a first sun gear set 210 and a second sun gear set 212. Each planetary gear 206 includes a first planetary gear set 214 and a second planetary gear set 216. The ring gear 208 includes a first ring gear set 218 and a second ring gear set 220.

As discussed in more detail below, the number of planetary gears may vary. In one embodiment, there are three planet gears 206. In another embodiment, there are two planet gears 206.

In the embodiment shown in fig. 4, the gear assembly 202 is a planetary gear configuration in which the ring gear 208 is substantially fixed (e.g., stationary) within the engine by a support structure 236. The sun gear 204 is driven by the input shaft (i.e., the low speed shaft 146). The planet carrier 222 is rotatably coupled to the plurality of planet gears 206, and the planet carrier 222 is configured to rotate in a circumferential direction 224 about the longitudinal centerline 120, which in turn drives a power output source 124 (e.g., a fan shaft), the power output source 124 being coupled to the planet carrier 222 and configured to rotate therewith to drive the fan assembly. In this embodiment, the low speed shaft 146 rotates in the same circumferential direction 226 as the direction 224 in which the fan shaft 124 rotates.

As shown in fig. 4, the sun gear 204 meshes with the planet gears 206 at first contact regions (i.e., sun gear mesh regions) 228, 230, and the ring gear 208 meshes with the planet gears 206 at second contact regions 232, 234 (i.e., ring gear mesh regions). The first contact regions 228, 230 are axially offset from the first contact regions 232, 234 such that the gear teeth of the sun gear are not axially aligned with the gear teeth of the ring gear relative to the longitudinal axis 120.

Referring again to fig. 4, the first and second ring gear sets 218, 220 are axially spaced from one another, and the first and second sun gear sets 210, 212 are positioned between the first and second ring gear sets 218, 220. As discussed in more detail below, the first contact region may not overlap the second contact region, or there may be a relatively small amount of overlap. Thus, in this embodiment, the ring gear 208 meshes with the planet gears 206 at an outward offset, and the sun gear 204 meshes with the planet gears 206 at an inward offset.

The sun gear and the ring gear may be axially offset such that at least 50% of the width of the sun gear mesh region does not axially overlap the ring gear mesh region. In other embodiments, the axial offset may be such that at least 25% of the width of the sun gear mesh region does not axially overlap the ring gear mesh region. Depending on the amount of offset, the axial offset of the first and second contact regions described herein may reduce and/or eliminate reverse bending of the planet gears. That is, due to the axial offset described herein, some or all of the planets do not alternately mesh with the sun and ring gears, which subjects the teeth to reverse bending stresses due to load reversals. The axial offset provided by this arrangement may provide a gear assembly with a higher gear ratio relative to a conventional single-stage epicyclic gear assembly designed to address the reverse bending limitation.

By reducing and/or eliminating reverse bending stresses on some or all of the planet gear teeth, the face width of the sun gear teeth may be reduced, which in turn may reduce the diameter of the sun gear. Because the gear ratio is related to the relative diameters of the ring gear and the sun gear, reducing the diameter of the sun gear while maintaining the diameter of the ring gear results in a corresponding increase in the gear ratio.

In some embodiments, the axial offset described herein may reduce the diameter of the sun gear by approximately 0.6 to 0.8 while maintaining a ring gear of the same diameter. Thus, for example, for a gear assembly having a planetary gear configuration with three planet gears, a sun gear diameter reduction of 0.7 may shift the gear ratio from 6: 1 to 8.2: 1, or a gear ratio from 6.5: 1 to 8.8: 1. this allows a single-stage gear assembly with a planetary gear configuration to achieve greater than or equal to 6: a gear ratio of 1, and in some embodiments, greater than or equal to 6.6: 1, greater than or equal to 7: 1, or greater than or equal to 8: 1. in other embodiments, the upper range of the gear assembly may be 14: 1, or in some cases an upper range of 12: 1.

FIG. 5 illustrates an example gear assembly 302 having an axially offset face width. Gear assembly 302 includes a diameter D_sSun gear 304 of diameter D_pA plurality of planet gears 306 and a diameter D_rRing gear 308. Each of the sun gear 304, planet gears 306, and ring gear 308 is a double helical gear having a first set of helical teeth and a second set of helical teeth that are inclined at an acute angle relative to each other.

The double helical sun gear 304 includes a first sun gear set 310 and a second sun gear set 312. Each of the double helical planet gears 306 includes a first planetary gear set 314 and a second planetary gear set 316. The double helical ring gear 308 includes a first ring gear set 318 and a second ring gear set 320. In some embodiments, there are two or three planet gears in the gear assembly.

In the embodiment shown in fig. 5, the gear assembly has a star gear configuration, wherein the planet carrier 322 is fixed (e.g., stationary) within the engine by a support structure 336. The sun gear 304 is driven by the input shaft (i.e., the low speed shaft 146). The ring gear 308 is configured to rotate in a direction 324 opposite a rotational direction 326 of the low speed shaft 146 to drive the power output source 124 (e.g., the fan shaft) and the fan assembly 104.

In a similar manner to fig. 4, the sun gear 304 meshes with the planet gears 306 at first contact regions 328, 330, and the ring gear 308 meshes with the planet gears 306 at second contact regions 332, 334, with the first contact regions 328, 330 being axially offset from the second contact regions 332, 334. As noted above, the axial offset provided by this arrangement may provide a gear assembly with a higher gear ratio relative to a conventional single-stage epicyclic gear assembly required to address the reverse bending stresses.

As described above, depending on the amount of offset, the axial offset of the first and second contact regions greatly reduces and/or eliminates the reverse bending of the planets, which in turn allows for a sun gear with a smaller diameter and a gear assembly with a high gear ratio. For example, for a gear assembly having a star gear configuration with three planet gears, reducing the diameter of the sun gear by 0.7 may result in a gear ratio from 5: 1 to 7.2: 1, and the gear ratio can be varied from 5.5: 1 increased to 7.8: 1. this allows a single stage gear assembly having a star gear configuration (such as a single stage gear assembly having a planetary gear configuration) to achieve a gear ratio of greater than or equal to 6: 1, and in some embodiments, greater than or equal to 6.6: 1, greater than or equal to 7: 1, and in some embodiments, greater than or equal to 8: 1. in other embodiments, the upper range of the gear assembly may be 14: 1, or in some cases an upper range of 12: 1.

fig. 6A, 6B, 7A and 7B are schematic illustrations of an exemplary gear assembly of the type shown in fig. 4 and 5. Some of the gear assembly structures (e.g., planet carrier, support structures) shown in fig. 6A-7B have been omitted for clarity. 6A-7B use the same reference numerals as provided for the planetary gear configuration in FIG. 4 for convenience, although it will be appreciated that the gear assembly shown in FIGS. 6A and 6B may be used with either the planetary gear configuration (FIG. 4) or the star gear configuration (FIG. 5).

Fig. 6A and 6B show a gear assembly having three planet gears 206, while fig. 7A and 7B show a gear assembly having two planet gears 206. Fig. 6B and 7B show the gear assembly of fig. 6A and 7A, respectively, with one of the double helical gear sets removed for clarity. As shown in these figures, the contact areas where the sun and planet gears mate are axially offset from the contact areas where the ring and planet gears mate to avoid and/or reduce the effect of reverse bending stresses on the planet gear teeth.

FIG. 8 illustrates another example gear assembly 402 having an axially offset face width. FIG. 8 has a planetary gear configuration and is similar to the gear assembly 202 shown in FIG. 4; however, in this embodiment, the ring gear meshes with the planet gears at an inward offset, and the sun gear meshes with the planet gears at an outward offset.

Referring to FIG. 8, the gear assembly 402 includes a diameter D_SSun gear 404 of diameter D_pA plurality of planet gears 406 and a diameter D_rRing gear 408. Each of the sun gear 404, planet gears 406, and ring gear 408 is a double helical gear having a first set of helical teeth and a second set of helical teeth that are inclined at an acute angle relative to each other. The sun gear 404 includes a first sun gear set 410 and a second sun gear set 412, the planet gears 406 include a first planetary gear set 414 and a second planetary gear set 416, and the ring gear 408 includes a first ring gear set 418 and a second ring gear set 420. The number of planet gears may vary as described elsewhere herein.

The embodiment shown in fig. 8 operates in the same manner as described above with respect to fig. 4, but with a different axial offset arrangement. As shown in fig. 8, the sun gear 404 meshes with the planet gears 406 at first contact regions 428, 430, and the ring gear 408 meshes with the planet gears 406 at second contact regions 432, 434, with the first contact regions 428, 430 being axially offset from the second contact regions 432, 434. The axial offset provided by this arrangement may provide a gear assembly with a higher gear ratio.

As in the embodiment of fig. 4, the axial offset of the first and second contact regions described herein greatly reduces and/or eliminates reverse bending of the planet gears, depending on the amount of offset, and provides for a reduction in sun gear diameter and an increase in gear ratio relative to conventional single stage gear assemblies, as described above.

FIG. 9 illustrates an example gear assembly 502 having an axially offset face width. FIG. 9 is a star gear configuration similar to the gear assembly 302 shown in FIG. 5; however, in this embodiment, the ring gear meshes with the planet gears at an inward offset, and the sun gear meshes with the planet gears at an outward offset.

Gear assembly 502 includes a diameter D_SSun gear 504 of diameter D_pA plurality of planet gears 506 and a diameter D_rRing gear 508. Each of the sun gear 504, planet gears 506, and ring gear 508 is a double helical gear having a first set of helical teeth and a second set of helical teeth that are inclined at an acute angle relative to each other. In particular, the sun gear 504 includes a first sun gear set 510 and a second sun gear set 512. Each planetary gear 506 includes a first planetary gear set 514 and a second planetary gear set 516. The ring gear 508 includes a first ring gear set 518 and a second ring gear set 520. The number of the planetary gears may be two or three.

In this embodiment, similar to that shown in fig. 5, the gear assembly has a star gear configuration, with the planet carrier 522 fixed (e.g., stationary) within the engine by a support structure 536. The sun gear 504 is driven by the input shaft (i.e., the low speed shaft 146). The ring gear 508 is configured to rotate in a direction 524 opposite a direction 526 of rotation of the low speed shaft 146 to drive the power output source 124 (e.g., a fan shaft) to drive the fan assembly.

In a manner similar to that shown in fig. 5, the sun gear 504 meshes with the planet gears 506 at first contact regions 528, 530, and the ring gear 508 meshes with the planet gears 506 at second contact regions 532, 534, and the first contact regions 528, 530 are axially offset from the second contact regions 532, 534 to provide a gear assembly with a higher gear ratio by reducing and/or eliminating reverse bending stresses on the planet gear teeth.

Fig. 10 provides an enlarged view of a portion of fig. 5. As shown in fig. 10, the first contact region 330 of the gear assembly 302 is axially spaced from the second contact region 334 such that there is no overlap between the first and second contact regions. An axial gap 340 is provided between the first contact region 330 and the second contact region 334. Similar gaps are provided between the other contact regions (i.e., the first contact region 328 and the second contact region 332). The gap 340 may be less than 25%, 20%, or 15% of the width 342 of the corresponding planetary gear set (e.g., planetary gear set 316). In some embodiments, the gap 340 between the first and second contact regions may be less than 10%, 5%, or 2% of the width 342 of the corresponding planetary gear set.

In non-overlapping embodiments, the width of a planetary gear set may be greater than the combined width of the respective sun and ring gear sets that mesh with the planetary gear sets. Thus, for example, the combined width of the first contact area 330 (i.e., the width of the second sun gear set 312) and the second contact area 334 (i.e., the width of the second ring gear set 320) is less than the width of the planet gears 342.

In other embodiments, the contact areas of the respective ring and sun gear sets may overlap. For example, fig. 11 shows an enlarged view of a portion of a gear assembly similar to that shown in fig. 5 but having overlapping contact areas.

Referring to fig. 11, the first contact region 330 of the gear assembly 302 is axially spaced from the second contact region 334; however, there is a relatively small amount of overlap 344 between the first contact region 330 and the second contact region 334. Similar overlap regions may be provided for the other contact regions (i.e., the first contact region 328 and the second contact region 332). The amount of overlap between the first and second contact regions (e.g., 330, 334) is preferably less than 15% of the width 342 of the respective planetary gear set (e.g., planetary gear set 316). More preferably, the amount of overlap 344 between the first and second contact regions is less than 10%, 5%, or 2% of the width 342 of the respective planetary gear set.

In overlapping embodiments, the width of a planetary gear set may be less than the combined width of the respective sun and ring gear sets that mesh with the planetary gear set. Thus, for example, in fig. 11, the combined width of the first contact area 330 (i.e., the width of the second sun gear set 312) and the second contact area 334 (i.e., the width of the second ring gear set 320) is greater than the width of the planet gears 342.

Although described above as a non-shrouded or open rotor engine in the above-described embodiments, it should be understood that the aspects of the disclosure provided herein may be applied to shrouded or ducted engines, partially ducted engines, aft fan engines or other turbine configurations, including configurations for marine, industrial or aviation propulsion systems. Certain aspects of the present disclosure may be applicable to turbofan, turboprop, or turboshaft engines. However, it should be appreciated that certain aspects of the present disclosure may address issues that may be specific to non-shrouded or developed rotary engines, such as, but not limited to, gear ratio, fan diameter, fan speed, length of engine (L), maximum diameter of engine (D)_max) L/D of engine_maxA desired cruising altitude and/or a desired operating cruising speed, or a combination thereof.

For example, fig. 12 is a cross-sectional schematic view of an exemplary embodiment of an engine 600, the engine 600 including a gear assembly 102 in combination with a ducted fan propulsion system. However, unlike the open rotor configuration of FIG. 2, fan assembly 104 and its fan blades 108 are contained within an annular fan casing 180, and vane assemblies 110 and vanes 112 extend radially between fan shroud 152 and an inner surface of fan casing 180. As described above, the gear assemblies disclosed herein may provide an increased gear ratio for a fixed gear envelope (e.g., a ring gear having the same size), or alternatively, a smaller diameter ring gear may be used to achieve the same gear ratio. Although fig. 12 shows an alternative ducted fan and an alternative fan duct (similar to that shown in fig. 2), it should be understood that such a gear assembly may be used with other turbofan engines (as well as other open rotor engines) that do not have any such structure.

This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.

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

1. a turbine engine, comprising: a fan assembly including a plurality of fan blades; a core engine including a turbine and an input shaft rotatable with the turbine; and a single-stage epicyclic gear assembly receiving the input shaft at a first speed and driving an output shaft coupled to the fan assembly at a second speed, the second speed being slower than the first speed. The gear assembly includes: a sun gear, a plurality of planet gears, and a ring gear, the sun gear rotating about a longitudinal centerline of the gear assembly; a sun gear mesh region along a longitudinal centerline of the gear assembly, the sun gear configured to contact the plurality of planet gears at the sun gear mesh region; a ring gear engagement region along a longitudinal centerline of the gear assembly, the ring gear configured to contact the plurality of planet gears at the ring gear engagement region, wherein the sun gear engagement region is axially offset from the ring gear engagement region along the longitudinal centerline such that at least 50% of the width of the sun gear engagement region does not axially overlap the ring gear engagement region.

2. The turbine engine of any item herein, wherein the sun gear, the plurality of planet gears, and the ring gear comprise double bevel gears, and the sun gear comprises a first sun gear set and a second sun gear set, each of the plurality of planet gears comprises a first planetary gear set and a second planetary gear set, and the ring gear comprises a first ring gear set and a second ring gear set.

3. The turbine engine of any item herein, wherein the first and second ring gear sets are axially spaced from each other along the longitudinal centerline, and the first and second sun gear sets are positioned between the first and second ring gear sets.

4. The turbine engine of any item herein, wherein the first and second sun gear sets are axially spaced from each other along the longitudinal centerline, and the first and second ring gear sets are positioned between the first and second sun gear sets.

5. A turbine engine according to any of the preceding claims, wherein the sun gear mesh zone comprises a first sun gear mesh zone at which the first sun gear set meshes with the first planetary gear set and a second sun gear mesh zone at which the second sun gear set meshes with the second planetary gear set, and the ring gear mesh zone comprises a first ring gear mesh zone at which the first ring gear set meshes with the first planetary gear set and a second ring gear mesh zone at which the second ring gear set meshes with the second planetary gear set.

6. A turbine engine in accordance with any of the clauses herein, wherein the first sun gear mesh region and the first ring gear mesh region do not axially overlap along the longitudinal centerline.

7. A turbine engine according to any item herein, wherein there is an axial gap between the first sun gear mesh region and the first ring gear mesh region.

8. A turbine engine according to any item herein, wherein a gap width of the axial gap is less than 15% of a width of the first planetary gear set, less than 10% of the width of the first planetary gear set, less than 5% of the width of the first planetary gear set, or less than 2% of the width of the first planetary gear set.

9. A turbine engine according to any of the clauses herein, wherein there is axial overlap between the first sun gear mesh region and the first ring gear mesh region, and the amount of axial overlap is less than 15% of the width of the first planetary gear set, less than 10% of the width of the first planetary gear set, less than 5% of the width of the first planetary gear set, or less than 2% of the width of the first planetary gear set.

10. The turbine of any clause herein, wherein the gear ratio of the gear assembly ranges from 6: 1 to 14: 1, from 6.6 to 12: 1, from 7: 1 to 12: 1 or from 8: 1 to 12: 1.

11. the turbine of any item herein, wherein the gear assembly is a planetary gear configuration in which the ring gear is fixed and does not rotate relative to the engine.

12. The turbine of any item herein, wherein the gear assembly is a star gear configuration in which the planet gears are fixed and do not rotate relative to the engine.

13. The turbomachine of any item herein, wherein the fan assembly is a single stage non-ducted fan blade.

14. The turbomachine of any item herein, wherein a width of the first planetary gear set is greater than a combined width of the first sun gear set and the first ring gear set.

15. The turbomachine of any item herein, wherein a width of the first planetary gear set is less than a combined width of the first sun gear set and the first ring gear set.

16. The turbine of any clause herein, wherein the fan assembly has ten to sixteen blades, or ten to fourteen blades, or twelve blades.

17. The turbine of any item herein, wherein a fan blade tip speed at cruise flight conditions is 650 to 900fps, or 700 to 800 fps.

18. The turbine of any item herein, wherein a Fan Pressure Ratio (FPR) of the fan assembly may be 1.04 to 1.10, or in some embodiments 1.05 to 1.08, as measured across the fan blades at cruise flight conditions.

19. A gear assembly that configures an input shaft at a first speed and drives an output shaft at a second speed slower than the first speed, the gear assembly comprising: a sun gear, a plurality of planet gears and a ring gear, the sun gear rotating about a longitudinal centerline of the gear assembly; a sun gear mesh region along a longitudinal centerline of the gear assembly, the sun gear configured to contact the plurality of planet gears at the sun gear mesh region; a ring gear engagement region along a longitudinal centerline of the gear assembly, the ring gear configured to contact the plurality of planet gears at the ring gear engagement region, wherein the sun gear engagement region is axially offset from the ring gear engagement region along the longitudinal centerline such that at least 50% of the width of the sun gear engagement region does not axially overlap the ring gear engagement region.

20. The gear assembly according to any one of the items herein, wherein the sun gear, the plurality of planet gears and the ring gear comprise a double bevel gear, and the sun gear comprises a first sun gear set and a second sun gear set, each of the plurality of planet gears comprises a first planetary gear set and a second planetary gear set, and the ring gear comprises a first ring gear set and a second ring gear set.

21. The gear assembly according to any one of the items herein, wherein the first and second ring gear sets are axially spaced from each other along the longitudinal centerline, and the first and second sun gear sets are positioned between the first and second ring gear sets.

22. The gear assembly according to any one of the items herein, wherein the first and second sun gear sets are axially spaced from each other along the longitudinal centerline, and the first and second ring gear sets are positioned between the first and second sun gear sets.

23. The gear assembly according to any one of the items herein, wherein the sun gear engagement zone includes a first sun gear engagement zone at which the first sun gear set meshes with the first planetary gear set and a second sun gear engagement zone at which the second sun gear set meshes with the second planetary gear set, and the ring gear engagement zone includes a first ring gear engagement zone at which the first ring gear set meshes with the first planetary gear set and a second ring gear engagement zone at which the second ring gear set meshes with the second planetary gear set.

24. A gear assembly according to any one of the clauses herein, wherein the first sun gear mesh region and the first ring gear mesh region do not axially overlap along the longitudinal centerline.

25. A gear assembly according to any one of the clauses herein, wherein there is an axial gap between the first sun gear mesh region and the first ring gear mesh region.

26. The gear assembly according to any of the items herein, wherein the gap width of the axial gap is less than 15% of the width of the first planetary gear set, less than 10% of the width of the first planetary gear set, less than 5% of the width of the first planetary gear set, or less than 2% of the width of the first planetary gear set.

27. The gear assembly according to any one of the items herein, wherein there is axial overlap between the first sun gear mesh region and the first ring gear mesh region, and an amount of axial overlap is less than 15% of a width of the first planetary gear set, less than 10% of the width of the first planetary gear set, less than 5% of the width of the first planetary gear set, or less than 2% of the width of the first planetary gear set.

28. A gear assembly according to any clause herein, wherein the gear ratio of the gear assembly ranges from 6: 1 to 14: 1, from 6.6 to 12: 1, from 7: 1 to 12: 1 or from 8: 1 to 12: 1.

29. a gear assembly according to any of the clauses herein, wherein the gear assembly is a planetary gear configuration or a star gear configuration.

30. The turbomachine of any item herein, wherein a width of the first planetary gear set is greater than a combined width of the first sun gear set and the first ring gear set, or wherein a width of the first planetary gear set is less than a combined width of the first sun gear set and the first ring gear set.

31. A turbine or gear assembly according to any clause herein, wherein the gear assembly is configured to provide an output torque ranging from 20kNm to 200kNm, or in other embodiments from 40kNm to 150 kNm.