Gearbox efficiency rating for turbomachine engines

A turbomachine engine can include a fan assembly, a vane assembly, a core engine, a gearbox, and a gearbox efficiency rating. The fan assembly can include a plurality of fan blades. The vane assembly can include a plurality of vanes, and the vanes can, in some instances, be disposed aft of the fan blades. The core engine can include 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 comprises a first rotational speed, the output is coupled to the fan assembly and has a second rotational speed, and a gear ratio of the first rotational speed to the second rotational speed is within a range of 4.1-14.0. The gearbox efficiency rating is 0.10-1.8.

CROSS-REFERENCE TO RELATED APPLICATION

This application claims the benefit of Italian Patent Application No. 102020000019171, filed Aug. 4, 2020, which is incorporated by reference herein.

FIELD

This disclosure relates generally to turbomachines including gearbox assemblies and, in particular, to apparatus and methods of determining gear assembly arrangements particular to certain turbomachine configurations.

BACKGROUND

A turbofan engine includes a core engine that drives a bypass fan. The bypass fan generates the majority of the thrust of the turbofan engine. The generated thrust can be used to move a payload (e.g., an aircraft).

In some instances, a turbofan engine is configured as a direct drive engine. Direct drive engines are configured such that a power turbine (e.g., a low-pressure turbine) of the core engine is directly coupled to the bypass fan. As such, the power turbine and the bypass fan rotate at the same rotational speed (i.e., the same rpm).

In other instances, a turbofan engine can be configured as a geared engine. Geared engines include a gearbox disposed between and interconnecting the bypass fan and power turbine of the core engine. The gearbox, for example, allows the power turbine of the core engine to rotate at a different speed than the bypass fan. Thus, the gearbox can, for example, allow the power turbine of the core engine and the bypass fan to operate at their respective rotational speeds for maximum efficiency and/or power production.

Despite certain advantages, geared turbofan engines can have one or more drawbacks. For example, including a gearbox in a turbofan engine introduces additional complexity to the engine. This can, for example, make engine development and/or manufacturing significantly more difficult. As such, there is a need for improved geared turbofan engines. There is also a need for devices and methods that can be used to develop and manufacture geared turbofan engines more efficiently and/or precisely.

BRIEF DESCRIPTION

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 through practice of the technology disclosed in the description.

Various turbomachine engines and gear assemblies are disclosed herein. The disclosed turbomachine engines comprise a gearbox. And the disclosed turbomachine engines are characterized or defined by a gearbox efficiency rating. The gearbox efficiency rating (GER) equals

Q⁡(D1.56T)1.53,
where Q is a gearbox oil flow rate an inlet of the gearbox measured in gallons per minute at a max takeoff condition, D is a diameter of the fan blades measured in inches, and T is a net thrust of the turbomachine engine measured in pounds force at the max takeoff condition. The gearbox efficiency rating may also be used, for example, to aid the development of the gearbox in relation to other engine parameters. The gearbox efficiency rating thus provides improved turbomachine engines and/or can help simplify one or more complexities of geared turbomachine engine development.

In particular embodiments, a turbomachine engine includes a fan assembly, a vane assembly, a core engine, a gearbox, and a gearbox efficiency rating. The fan assembly includes a plurality of fan blades. The vane assembly includes a plurality of vanes. 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 comprises a first rotational speed, the output is coupled to the fan assembly and has a second rotational speed, and a gear ratio of the first rotational speed to the second rotational speed is within a range of 4.1-14.0. The gearbox efficiency rating is 0.10-1.8.

These and other features, aspects, and/or advantages of the present disclosure will become better understood with reference to the following description and the 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.

DETAILED DESCRIPTION

One or more components of the turbomachine engine or gear assembly described hereinbelow may be manufactured or formed using any suitable process, such as an additive manufacturing process, such as a 3-D printing process. The use of such a process may allow such component to be formed integrally, as a single monolithic component, or as any suitable number of sub-components. In particular, the additive manufacturing process may allow such component to be integrally formed and include a variety of features not possible when using prior manufacturing methods. For example, the additive manufacturing methods described herein enable the manufacture of heat exchangers having unique features, configurations, thicknesses, materials, densities, fluid passageways, headers, and mounting structures that may not have been possible or practical using prior manufacturing methods. Some of these features are described herein.

Referring now to the drawings,FIG. 1is an exemplary embodiment of an engine100including a gear assembly102according to aspects of the present disclosure. The engine100includes a fan assembly104driven by a core engine106. In various embodiments, the core engine106is a Brayton cycle system configured to drive the fan assembly104. The core engine106is shrouded, at least in part, by an outer casing114. The fan assembly104includes a plurality of fan blades108. A vane assembly110extends from the outer casing114in a cantilevered manner. Thus, the vane assembly110can also be referred to as an unducted vane assembly. The vane assembly110, including a plurality of vanes112, is positioned in operable arrangement with the fan blades108to provide thrust, control thrust vector, abate or re-direct undesired acoustic noise, and/or otherwise desirably alter a flow of air relative to the fan blades108.

In some embodiments, the fan assembly104includes eight (8) to twenty (20) fan blades108. In particular embodiments, the fan assembly104includes ten (10) to eighteen (18) fan blades108. In certain embodiments, the fan assembly104includes twelve (12) to sixteen (16) fan blades108. In some embodiments, the vane assembly110includes three (3) to thirty (30) vanes112. In certain embodiments, the vane assembly110includes an equal or fewer quantity of vanes112to fan blades108. For example, in particular embodiments, the engine100includes twelve (12) fan blades108and ten (10) vanes112. In other embodiments, the vane assembly110includes a greater quantity of vanes112to fan blades108. For example, in particular embodiments, the engine100includes ten (10) fan blades108and twenty-three (23) vanes112.

In certain embodiments, such as depicted inFIG. 1, the vane assembly110is positioned downstream or aft of the fan assembly104. However, it should be appreciated that in some embodiments, the vane assembly110may be positioned upstream or forward of the fan assembly104. In still various embodiments, the engine100may include a first vane assembly positioned forward of the fan assembly104and a second vane assembly positioned aft of the fan assembly104. The fan assembly104may be configured to desirably adjust pitch at one or more fan blades108, such as to control thrust vector, abate or re-direct noise, and/or alter thrust output. The vane assembly110may be configured to desirably adjust pitch at one or more vanes112, such as to control thrust vector, abate or re-direct noise, and/or alter thrust output. Pitch control mechanisms at one or both of the fan assembly104or the vane assembly110may co-operate to produce one or more desired effects described above.

In certain embodiments, such as depicted inFIG. 1, the engine100is an un-ducted thrust producing system, such that the plurality of fan blades108is unshrouded by a nacelle or fan casing. As such, in various embodiments, the engine100may be configured as an unshrouded turbofan engine, an open rotor engine, or a propfan engine. In particular embodiments, the engine100is an unducted rotor engine with a single row of fan blades108. The fan blades108can have a large diameter, such as may be suitable for high bypass ratios, high cruise speeds (e.g., comparable to aircraft with turbofan engines, or generally higher cruise speed than aircraft with turboprop engines), high cruise altitude (e.g., comparable to aircraft with turbofan engines, or generally higher cruise speed than aircraft with turboprop engines), and/or relatively low rotational speeds.

The fan blades108comprise a diameter (Dfan). It should be noted that for purposes of illustration only half of the Dfanis shown (i.e., the radius of the fan). In some embodiments, the Dfanis 72-216 inches. In particular embodiments the Dfanis 100-200 inches. In certain embodiments, the Dfanis 120-190 inches. In other embodiments, the Dfanis 72-120 inches. In yet other embodiments, the Dfanis 50-80 inches.

In some embodiments, the fan blade tip speed at a cruise flight condition can be 650 to 900 fps, or 700 to 800 fps. A fan pressure ratio (FPR) for the fan assembly104can be 1.04 to 1.10, or in some embodiments 1.05 to 1.08, as measured across the fan blades at a cruise flight condition.

Cruise altitude is generally an altitude at which an aircraft levels after climb and prior to descending to an approach flight phase. In various embodiments, the engine is applied to a vehicle with a cruise altitude up to approximately 65,000 ft. In certain embodiments, cruise altitude is between approximately 28,000 ft. and approximately 45,000 ft. In still certain embodiments, cruise altitude is expressed in flight levels (FL) based on a standard air pressure at sea level, in which a cruise flight condition is between FL280 and FL650. In another embodiment, cruise flight condition is between FL280 and FL450. In still certain embodiments, cruise altitude is defined based at least on a barometric pressure, in which cruise altitude is between approximately 4.85 psia and approximately 0.82 psia based on a sea-level pressure of approximately 14.70 psia and sea-level temperature at approximately 59 degrees Fahrenheit. In another embodiment, cruise altitude is between approximately 4.85 psia and approximately 2.14 psia. It should be appreciated that in certain embodiments, the ranges of cruise altitude defined by pressure may be adjusted based on a different reference sea-level pressure and/or sea-level temperature.

The core engine106is generally encased in outer casing114defining one half of a core diameter (Dcore), which may be thought of as the maximum extent from the centerline axis (datum for R). In certain embodiments, the engine100includes a length (L) from a longitudinally (or axial) forward end116to a longitudinally aft end118. In various embodiments, the engine100defines a ratio of L/Dcorethat provides for reduced installed drag. In one embodiment, L/Dcoreis at least 2. In another embodiment, L/Dcoreis at least 2.5. In some embodiments, the L/Dcore, is less than 5, less than 4, and less than 3. In various embodiments, it should be appreciated that the L/Dcoreis for a single unducted rotor engine.

The reduced installed drag may further provide for improved efficiency, such as improved specific fuel consumption. Additionally, or alternatively, the reduced drag may provide for cruise altitude engine and aircraft operation at or above Mach 0.5. In certain embodiments, the L/Dcore, the fan assembly104, and/or the vane assembly110separately or together configure, at least in part, the engine100to 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.

Referring still toFIG. 1, the core engine106extends in a radial direction (R) relative to an engine centerline axis120. The gear assembly102receives power or torque from the core engine106through a power input source122and provides power or torque to drive the fan assembly104, in a circumferential direction C about the engine centerline axis120, through a power output source124.

The gear assembly102of the engine100can include a plurality of gears, including an input and an output. The gear assembly can also include one or more intermediate gears disposed between and/or interconnecting the input and the output. The input can be coupled to a turbine section of the core engine106and can comprise a first rotational speed. The output can be coupled to the fan assembly and can have a second rotational speed. In some embodiments, a gear ratio of the first rotational speed to the second rotational speed is greater than 4.1 (e.g., within a range of 4.1-14.0).

The gear assembly102(which can also be referred to as “a gearbox”) can comprise various types and/or configuration. For example, in some embodiments, the gearbox is an epicyclic gearbox configured in a star gear configuration. Star gear configurations comprise a sun gear, a plurality of star gears (which can also be referred to as “planet gears”), and a ring gear. The sun gear is the input and is coupled to the power turbine (e.g., the low-pressure turbine) such that the sun gear and the power turbine rotate at the same rotational speed. The star gears are disposed between and interconnect the sun gear and the ring gear. The star gears are rotatably coupled to a fixed carrier. As such, the star gears can rotate about their respective axes but cannot collectively orbit relative to the sun gear or the ring gear. As another example, the gearbox is an epicyclic gearbox configured in a planet gear configuration. Planet gear configurations comprise a sun gear, a plurality of planet gears, and a ring gear. The sun gear is the input and is coupled to the power turbine. The planet gears are disposed between and interconnect the sun gear and the ring gear. The planet gears are rotatably coupled to a rotatable carrier. As such, the planet gears can rotate about their respective axes and also collectively rotate together with the carrier relative to the sun gear and the ring gear. The carrier is the output and is coupled to the fan assembly. The ring gear is fixed from rotation.

In some embodiments, the gearbox is a single-stage gearbox (e.g.,FIGS. 10-11). In other embodiments, the gearbox is a multi-stage gearbox (e.g.,FIGS. 9 and 12). In some embodiments, the gearbox is an epicyclic gearbox. In some embodiments, the gearbox is a non-epicyclic gearbox (e.g., a compound gearbox—FIG. 13).

As noted above, the gear assembly can be used to reduce the rotational speed of the output relative to the input. In some embodiments, a gear ratio of the input rotational speed to the output rotational speed is greater than 4.1. For example, in particular embodiments, the gear ratio is within a range of 4.1-14.0, within a range of 4.5-14.0, or within a range of 6.0-14.0. In certain embodiments, the gear ratio is within a range of 4.5-12 or within a range of 6.0-11.0. As such, in some embodiments, the fan assembly can be configured to rotate at a rotational speed of 700-1500 rpm at a cruise flight condition, while the power turbine (e.g., the low-pressure turbine) is configured to rotate at a rotational speed of 2,500-15,000 rpm at a cruise flight condition. In particular embodiments, the fan assembly can be configured to rotate at a rotational speed of 850-1350 rpm at a cruise flight condition, while the power turbine is configured to rotate at a rotational speed of 5,000-10,000 rpm at a cruise flight condition.

Various gear assembly configurations are depicted schematically inFIGS. 9-13. These gearboxes can be used any of the engines disclosed herein, including the engine100. Additional details regarding the gearboxes are provided below.

FIG. 2shows a cross-sectional view of an engine200, which is configured as an exemplary embodiment of an open rotor propulsion engine. The engine200is generally similar to the engine100and corresponding components have been numbered similarly. For example, the gear assembly of the engine100is numbered “102” and the gear assembly of the engine200is numbered “202,” and so forth. In addition to the gear assembly202, the engine200comprises a fan assembly204that includes a plurality of fan blades208distributed around the engine centerline axis220. Fan blades208are circumferentially arranged in an equally spaced relation around the engine centerline axis220, and each fan blade208has a root225and a tip226, and an axial span defined therebetween, as well as a central blade axis228.

The core engine206includes a compressor section230, a combustion section232, and a turbine section234(which may be referred to as “an expansion section”) together in a serial flow arrangement. The core engine206extends circumferentially relative to an engine centerline axis220. The core engine206includes a high-speed spool that includes a high-speed compressor236and a high-speed turbine238operably rotatably coupled together by a high-speed shaft240. The combustion section232is positioned between the high-speed compressor236and the high-speed turbine238.

The combustion section232may be configured as a deflagrative combustion section, a rotating detonation combustion section, a pulse detonation combustion section, and/or other appropriate heat addition system. The combustion section232may be configured as one or more of a rich-burn system or a lean-burn system, or combinations thereof. In still various embodiments, the combustion section232includes an annular combustor, a can combustor, a cannular combustor, a trapped vortex combustor (TVC), or other appropriate combustion system, or combinations thereof.

The core engine206also includes a booster or low-pressure compressor positioned in flow relationship with the high-pressure compressor236. The low-pressure compressor242is rotatably coupled with the low-pressure turbine244via a low-speed shaft246to enable the low-pressure turbine244to drive the low-pressure compressor242. The low-speed shaft246is also operably connected to the gear assembly202to provide power to the fan assembly204, such as described further herein.

It should be appreciated that the terms “low” and “high,” or their respective comparative degrees (e.g., “lower” and “higher”, where applicable), when used with compressor, turbine, shaft, or spool components, each refer to relative pressures and/or relative speeds within an engine unless otherwise specified. For example, a “low spool” or “low-speed shaft” defines a component configured to operate at a rotational speed, such as a maximum allowable rotational speed, lower than a “high spool” or “high-speed shaft” of the engine. Alternatively, unless otherwise specified, the aforementioned terms may be understood in their superlative degree. For example, a “low turbine” or “low-speed turbine” may refer to the lowest maximum rotational speed turbine within a turbine section, a “low compressor” or “low speed compressor” may refer to the lowest maximum rotational speed turbine within a 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, the low-speed spool refers to a lower maximum rotational speed than the high-speed spool. It should further be appreciated that the terms “low” or “high” in such aforementioned regards may additionally, or alternatively, be understood as relative to minimum allowable speeds, or minimum or maximum allowable speeds relative to normal, desired, steady state, etc. operation of the engine.

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

In some embodiments, a low-pressure turbine is a counter-rotating low-pressure turbine comprising inner blade stages and outer blade stages. The inner blade stages extend radially outwardly from an inner shaft, and the outer blade stages extend radially inwardly from an outer drum. In particular embodiments, the counter-rotating low-pressure turbine comprises three inner blade stages and three outer blade stages, which can collectively be referred to as a six stage low-pressure turbine. In other embodiments, the counter-rotating low-pressure turbine comprises four inner blade stages and three outer blade stages, which can be collectively be referred to as a seven stage low-pressure turbine.

As discussed in more detail below, the core engine206includes the gear assembly202that is configured to transfer power from the turbine section234and reduce an output rotational speed at the fan assembly204relative to the low-speed turbine244. Embodiments of the gear assembly202depicted and described herein can allow for gear ratios suitable for large-diameter unducted fans (e.g., gear ratios of 4.1-14.0, 4.5-14.0, and/or 6.0-14.0). Additionally, embodiments of the gear assembly202provided herein may be suitable within the radial or diametrical constraints of the core engine206within the outer casing214.

Various gearbox configurations are depicted schematically inFIGS. 9-13. These gearboxes can be used in any of the engines disclosed herein, including the engine200. Additional details regarding the gearboxes are provided below.

Engine200also includes a vane assembly210comprising a plurality of vanes212disposed around engine centerline axis220. Each vane212has a root248and a tip250, and a span defined therebetween. Vanes212can be arranged in a variety of manners. In some embodiments, for example, they are not all equidistant from the rotating assembly.

In some embodiments, vanes212are mounted to a stationary frame and do not rotate relative to the engine centerline axis220, but may include a mechanism for adjusting their orientation relative to their axis254and/or relative to the fan blades208. For reference purposes,FIG. 2depicts a forward direction denoted with arrow F, which in turn defines the forward and aft portions of the system.

As depicted inFIG. 2, the fan assembly204is located forward of the core engine106with the exhaust256located aft of core engine206in a “puller” configuration. Other configurations are possible and contemplated as within the scope of the present disclosure, such as what may be termed a “pusher” configuration embodiment where the engine core is located forward of the fan assembly. The selection of “puller” or “pusher” configurations may be made in concert with the selection of mounting orientations with respect to the airframe of the intended aircraft application, and some may be structurally or operationally advantageous depending upon whether the mounting location and orientation are wing-mounted, fuselage-mounted, or tail-mounted configurations.

Left- or right-handed engine configurations, useful for certain installations in reducing the impact of multi-engine torque upon an aircraft, can be achieved by mirroring the airfoils (e.g.,208,212) such that the fan assembly204rotates clockwise for one propulsion system and counterclockwise for the other propulsion system. Alternatively, an optional reversing gearbox can be provided to permit a common gas turbine core and low-pressure turbine to be used to rotate the fan blades either clockwise or counterclockwise, i.e., to provide either left- or right-handed configurations, as desired, such as to provide a pair of oppositely-rotating engine assemblies can be provided for certain aircraft installations while eliminating the need to have internal engine parts designed for opposite rotation directions.

The engine200also includes the gear assembly202which includes a gear set for decreasing the rotational speed of the fan assembly204relative to the low-speed turbine244. In operation, the rotating fan blades208are driven by the low-speed turbine244via gear assembly202such that the fan blades208rotate around the engine centerline axis220and generate thrust to propel the engine200, and hence an aircraft on which it is mounted, in the forward direction F.

In some embodiments, a gear ratio of the input rotational speed to the output rotational speed is greater than 4.1. For example, in particular embodiments, the gear ratio is within a range of 4.1-14.0, within a range of 4.5-14.0, or within a range of 6.0-14.0. In certain embodiments, the gear ratio is within a range of 4.5-12 or within a range of 6.0-11.0. As such, in some embodiments, the fan assembly can be configured to rotate at a rotational speed of 700-1500 rpm at a cruise flight condition, while the power turbine (e.g., the low-pressure turbine) is configured to rotate at a rotational speed of 5,000-10,000 rpm at a cruise flight condition. In particular embodiments, the fan assembly can be configured to rotate at a rotational speed of 850-1350 rpm at a cruise flight condition, while the power turbine is configured to rotate at a rotational speed of 5,500-9,500 rpm a cruise flight condition.

It may be desirable that either or both of the fan blades208or the vanes212to incorporate a pitch change mechanism such that the blades can be rotated with respect to an axis of pitch rotation (annotated as228and254, respectively) either independently or in conjunction with one another. Such pitch change can be utilized to vary thrust and/or swirl effects under various operating conditions, including to provide a thrust reversing feature which may be useful in certain operating conditions such as upon landing an aircraft.

Vanes212can be sized, shaped, and configured to impart a counteracting swirl to the fluid so that in a downstream direction aft of both fan blades208and vanes212the fluid has a greatly reduced degree of swirl, which translates to an increased level of induced efficiency. Vanes212may have a shorter span than fan blades208, as shown inFIG. 2. For example, vanes212may have a span that is at least 50% of a span of fan blades208. In some embodiments, the span of the vanes can be the same or longer than the span as fan blades208, if desired. Vanes212may be attached to an aircraft structure associated with the engine200, as shown inFIG. 2, or another aircraft structure such as a wing, pylon, or fuselage. Vanes212may be fewer or greater in number than, or the same in number as, the number of fan blades208. In some embodiments, the number of vanes212are greater than two, or greater than four, in number. Fan blades208may be sized, shaped, and contoured with the desired blade loading in mind.

In the embodiment shown inFIG. 2, an annular 360-degree inlet258is located between the fan assembly204and the vane assembly210, and provides a path for incoming atmospheric air to enter the core engine206radially inwardly of at least a portion of the vane assembly210. Such a location may be advantageous for a variety of reasons, including management of icing performance as well as protecting the inlet258from various objects and materials as may be encountered in operation.

In the exemplary embodiment ofFIG. 2, in addition to the open rotor or unducted fan assembly204with its plurality of fan blades208, an optional ducted fan assembly260is included behind fan assembly204, such that the engine200includes both a ducted and an unducted fan which both serve to generate thrust through the movement of air at atmospheric temperature without passage through the core engine206. The ducted fan assembly260is shown at about the same axial location as the vane212, and radially inward of the root248of the vane212. Alternatively, the ducted fan assembly260may be between the vane212and core duct262, or be farther forward of the vane212. The ducted fan assembly260may be driven by the low-pressure turbine244, or by any other suitable source of rotation, and may serve as the first stage of the low-pressure compressor242or may be operated separately. Air entering the inlet258flows through an inlet duct264and then is divided such that a portion flows through a core duct262and a portion flows through a fan duct266. Fan duct266may incorporate heat exchangers268and exhausts to the atmosphere through an independent fixed or variable nozzle270aft of the vane assembly210, at the aft end of the fan cowl252and outside of the engine core cowl272. Air flowing through the fan duct266thus “bypasses” the core of the engine and does not pass through the core.

Thus, in the exemplary embodiment, engine200includes an unducted fan formed by the fan blades208, followed by the ducted fan assembly260, which directs airflow into two concentric or non-concentric ducts262and266, thereby forming a three-stream engine architecture with three paths for air which passes through the fan assembly204.

In the exemplary embodiment shown inFIG. 2, a slidable, moveable, and/or translatable plug nozzle274with an actuator may be included in order to vary the exit area of the nozzle270. A plug nozzle is typically an annular, symmetrical device which regulates the open area of an exit such as a fan stream or core stream by axial movement of the nozzle such that the gap between the nozzle surface and a stationary structure, such as adjacent walls of a duct, varies in a scheduled fashion thereby reducing or increasing a space for airflow through the duct. Other suitable nozzle designs may be employed as well, including those incorporating thrust reversing functionality. Such an adjustable, moveable nozzle may be designed to operate in concert with other systems such as VBV's, VSV's, or blade pitch mechanisms and may be designed with failure modes such as fully-open, fully-closed, or intermediate positions, so that the nozzle270has a consistent “home” position to which it returns in the event of any system failure, which may prevent commands from reaching the nozzle270and/or its actuator.

In some embodiments, a mixing device276can be included in a region aft of a core nozzle278to aid in mixing the fan stream and the core stream to improve acoustic performance by directing core stream outward and fan stream inward.

Since the engine200shown inFIG. 2includes both an open rotor fan assembly204and a ducted fan assembly260, the thrust output of both and the work split between them can be tailored to achieve specific thrust, fuel burn, thermal management, and/or acoustic signature objectives which may be superior to those of a typical ducted fan gas turbine propulsion assembly of comparable thrust class. The ducted fan assembly260, by lessening the proportion of the thrust required to be provided by the unducted fan assembly104, may permit a reduction in the overall fan diameter of the unducted fan assembly and thereby provide for installation flexibility and reduced weight.

Operationally, the engine200may include a control system that manages the loading of the respective open and ducted fans, as well as potentially 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 operational conditions associated with aircraft operation. For example, in climb mode the ducted fan may operate at maximum pressure ratio there-by maximizing the thrust capability of stream, while in cruise mode, the ducted fan may operate a lower pressure ratio, raising overall efficiency through reliance on thrust from the unducted fan. Nozzle actuation modulates the ducted fan operating line and overall engine fan pressure ratio independent of total engine airflow.

The ducted fan stream flowing through fan duct266may include one or more heat exchangers268for removing heat from various fluids used in engine operation (such as an air-cooled oil cooler (ACOC), cooled cooling air (CCA), etc.). The heat exchangers268may take advantage of the integration into the fan duct266with reduced performance penalties (such as fuel efficiency and thrust) compared with traditional ducted fan architectures, due to not impacting the primary source of thrust which is, in this case, the unducted fan stream. Heat exchangers may cool fluids such as gearbox oil, engine sump oil, thermal transport fluids such as supercritical fluids or commercially available single-phase or two-phase fluids (supercritical CO2, EGV, Slither 800, liquid metals, etc.), engine bleed air, etc. Heat exchangers may also be made up of different segments or passages that cool different working fluids, such as an ACOC paired with a fuel cooler. Heat exchangers268may be incorporated into a thermal management system which provides for thermal transport via a heat exchange fluid flowing through a network to remove heat from a source and transport it to a heat exchanger.

Since the fan pressure ratio is higher for the ducted fan than for the unducted fan, the fan duct provides an environment where more compact heat exchangers may be utilized than would be possible if installed on the outside of the core cowl in the unducted fan stream. Fan bypass air is at a very low fan pressure ratio (FPR) (1.05 to 1.08), making it difficult to drive air through heat exchangers. Without the availability of a fan duct as described herein, scoops or booster bleed air may be required to provide cooling air to and through heat exchangers. A set of parameters can be developed around heat exchangers in the fan duct, based on heat load, heat exchanger size, ducted fan stream corrected flow, and ducted fan stream temperature.

The fan duct266also provides other advantages in terms of reduced nacelle drag, enabling a more aggressive nacelle close-out, improved core stream particle separation, and inclement weather operation. By exhausting the fan duct flow over the core cowl, this aids in energizing the boundary layer and enabling the option of a steeper nacelle close out angle between the maximum dimension of the engine core cowl272and the exhaust256. The close-out angle is normally limited by air flow separation, but boundary layer energization by air from the fan duct266exhausting over the core cowl reduces air flow separation. This yields a shorter, lighter structure with less frictional surface drag.

The fan assembly and/or vane assembly can be shrouded or unshrouded (as shown inFIGS. 1 and 2). Although not shown, an optional annular shroud or duct can be coupled to the vane assembly210and located distally from the engine centerline axis220relative to the vanes212. In addition to the noise reduction benefit, the duct may provide improved vibratory response and structural integrity of the vanes212by coupling them into an assembly forming an annular ring or one or more circumferential sectors, i.e., segments forming portions of an annular ring linking two or more of the vanes212. The duct may also allow the pitch of the vanes to be varied more easily. For example,FIGS. 3-4, discussed in more detail below, disclose embodiments in which both the fan assembly and vane assembly are shrouded.

Although depicted above as an unshrouded or open rotor engine in the embodiments depicted above, it should be appreciated that aspects of the disclosure provided herein may be applied to shrouded or ducted engines, partially ducted engines, aft-fan engines, or other turbomachine configurations, including those for marine, industrial, or aero-propulsion systems. Certain aspects of the disclosure may be applicable to turbofan, turboprop, or turboshaft engines. However, it should be appreciated that certain aspects of the disclosure may address issues that may be particular to unshrouded or open rotor engines, such as, but not limited to, issues related to gear ratios, fan diameter, fan speed, length (L) of the engine, maximum diameter of the core engine (Dcore) of the engine, L/Dcore of the engine, desired cruise altitude, and/or desired operating cruise speed, or combinations thereof.

FIG. 3is a schematic cross-sectional view of a gas turbine engine in accordance with an exemplary embodiment of the present disclosure. More particularly, for the embodiment ofFIG. 3, the gas turbine engine is a high-bypass turbofan jet engine300, referred to herein as “turbofan engine300.” As shown inFIG. 3, the turbofan engine300defines an axial direction A (extending parallel to a longitudinal centerline302provided for reference) and a radial direction R (extending perpendicular to the axial direction A). In general, the turbofan300includes a fan section304and a core engine306disposed downstream from the fan section304. The engine300also includes a gear assembly or power gear box336having a plurality of gears for coupling a gas turbine shaft to a fan shaft. The position of the power gear box336is not limited to that as shown in the exemplary embodiment of turbofan300. For example, the position of the power gear box336may vary along the axial direction A.

The exemplary core engine306depicted generally includes a substantially tubular outer casing308that defines an annular inlet310. The outer casing308encases, in serial flow relationship, a compressor section including a booster or low-pressure (LP) compressor312and a high-pressure (HP) compressor314; a combustion section316; a turbine section including a high-pressure (HP) turbine318and a low-pressure (LP) turbine320; and a jet exhaust nozzle section322. A high-pressure (HP) shaft or spool324drivingly connects the HP turbine318to the HP compressor314. A low-pressure (LP) shaft or spool326drivingly connects the LP turbine320to the LP compressor312. Additionally, the compressor section, combustion section316, and turbine section together define at least in part a core air flowpath327extending therethrough.

A gear assembly of the present disclosure is compatible with standard fans, variable pitch fans, or other configurations. For the embodiment depicted, the fan section304may include a variable pitch fan328having a plurality of fan blades330coupled to a disk332in a spaced apart manner. As depicted, the fan blades330extend outwardly from disk332generally along the radial direction R. Each fan blade330is rotatable relative to the disk332about a pitch axis P by virtue of the fan blades330being operatively coupled to a suitable actuation member334configured to collectively vary the pitch of the fan blades330. The fan blades330, disk332, and actuation member334are together rotatable about the longitudinal axis302by LP shaft326across a gear assembly or power gear box336. A gear assembly336may enable a speed change between a first shaft, e.g., LP shaft326, and a second shaft, e.g., LP compressor shaft and/or fan shaft. For example, in one embodiment, the gear assembly336may be disposed in an arrangement between a first shaft and a second shaft such as to reduce an output speed from one shaft to another shaft.

More generally, the gear assembly336can be placed anywhere along the axial direction A to decouple the speed of two shafts, whenever it is convenient to do so from a component efficiency point of view, e.g., faster LP turbine and slower fan and LP compressor or faster LP turbine and LP compressor and slower fan.

Referring still to the exemplary embodiment ofFIG. 3, the disk332is covered by rotatable front nacelle338aerodynamically contoured to promote an airflow through the plurality of fan blades330. Additionally, the exemplary fan section304includes an annular fan casing or outer nacelle340that circumferentially surrounds the fan328and/or at least a portion of the core engine306. The nacelle340is, for the embodiment depicted, supported relative to the core engine306by a plurality of circumferentially-spaced outlet guide vanes342. Additionally, a downstream section344of the nacelle340extends over an outer portion of the core engine306so as to define a bypass airflow passage346therebetween.

During operation of the turbofan engine300, a volume of air348enters the turbofan300through an associated inlet350of the nacelle340and/or fan section304. As the volume of air348passes across the fan blades330, a first portion of the air348as indicated by arrows352is directed or routed into the bypass airflow passage346and a second portion of the air348as indicated by arrow354is directed or routed into the LP compressor312. The ratio between the first portion of air352and the second portion of air354is commonly known as a bypass ratio. The pressure of the second portion of air354is then increased as it is routed through the high-pressure (HP) compressor314and into the combustion section316, where it is mixed with fuel and burned to provide combustion gases356.

The combustion gases356are routed through the HP turbine318where a portion of thermal and/or kinetic energy from the combustion gases356is extracted via sequential stages of HP turbine stator vanes358that are coupled to the outer casing308and HP turbine rotor blades360that are coupled to the HP shaft or spool324, thus causing the HP shaft or spool324to rotate, thereby supporting operation of the HP compressor314. The combustion gases356are then routed through the LP turbine320where a second portion of thermal and kinetic energy is extracted from the combustion gases356via sequential stages of LP turbine stator vanes362that are coupled to the outer casing308and LP turbine rotor blades364that are coupled to the LP shaft or spool326, thus causing the LP shaft or spool326to rotate, thereby supporting operation of the LP compressor312and/or rotation of the fan328.

The combustion gases356are subsequently routed through the jet exhaust nozzle section322of the core engine306to provide propulsive thrust. Simultaneously, the pressure of the first portion of air352is substantially increased as the first portion of air352is routed through the bypass airflow passage346before it is exhausted from a fan nozzle exhaust section366of the turbofan300, also providing propulsive thrust. The HP turbine318, the LP turbine320, and the jet exhaust nozzle section322at least partially define a hot gas path368for routing the combustion gases356through the core engine306.

For example,FIG. 4is a cross-sectional schematic illustration of an exemplary embodiment of an engine400that includes a gear assembly402in combination with a ducted fan assembly404and a core engine406. However, unlike the open rotor configuration of the engine200, the fan assembly404and its fan blades408are contained within an annular fan case480(which can also be referred to as “a nacelle”) and the vane assembly410and the vanes412extend radially between the fan cowl452(and/or the engine core cowl472) and the inner surface of the fan case480. As discussed above, the gear assemblies disclosed herein can provide for increased gear ratios for a fixed gear envelope (e.g., with the same size ring gear), or alternatively, a smaller diameter ring gear may be used to achieve the same gear ratios.

The core engine400comprises a compressor section430, a combustor section432, and a turbine section434. The compressor section430can include a high-pressure compressor436and a booster or a low-pressure compressor442. The turbine section434can include a high-pressure turbine438and a low-pressure turbine444. The low-pressure compressor442is positioned forward of and in flow relationship with the high-pressure compressor436. The low-pressure compressor442is rotatably coupled with the low-pressure turbine444via a low-speed shaft446to enable the low-pressure turbine444to drive the low-pressure compressor442(and a ducted fan460). The low-speed shaft446is also operably connected to the gear assembly402to provide power to the fan assembly404. The high-pressure compressor436is rotatably coupled with the high-pressure turbine438via a high-speed shaft440to enable the high-pressure turbine438to drive the high-pressure compressor436.

In some embodiments, the engine400can comprise a pitch change mechanism482coupled to the fan assembly404and configured to vary the pitch of the fan blades408. In certain embodiments, the pitch change mechanism482can be a linear actuated pitch change mechanism.

In some embodiments, the engine400can comprise a variable fan nozzle. Operationally, the engine400may include a control system that manages the loading of the fan, as well as potentially 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 operational conditions associated with aircraft operation. For example, nozzle actuation modulates the fan operating line and overall engine fan pressure ratio independent of total engine airflow.

In some embodiments, an engine (e.g., the engine100, the engine200, and/or the engine400) can comprise a counter-rotating low-pressure turbine. For example,FIGS. 5-6depict schematic cross-sectional illustrations of counter-rotating low-pressure turbines. In particular,FIG. 5depicts a counter-rotating turbine500, andFIG. 6depicts a counter-rotating turbine600. The counter-rotating turbines comprise inner blade stages and outer blade stages arranged in an alternating inner-outer configuration. In other words, the counter-rotating turbines do not comprise stator vanes disposed between the blade stages.

Referring toFIG. 5, the counter-rotating turbine500comprises a plurality of inner blade stages502and a plurality of outer blade stages504. More specifically, the counter-rotating turbine500includes three inner blades stages502that are coupled to and extend radially outwardly from an inner shaft506(which can also be referred to as “a rotor”) and three outer blade stages504that are coupled to extend radially inwardly from an outer shaft508(which can also be referred to as “a drum”). In this manner, the counter-rotating turbine500can be considered a six stage turbine.

Referring toFIG. 6, the counter-rotating turbine600comprises a plurality of inner blade stages602and a plurality of outer blade stages604. More specifically, the counter-rotating turbine600includes four inner blades stages602that are coupled to and extend radially outwardly from an inner shaft606and three outer blade stages604that are coupled to extend radially inwardly from an outer shaft608. In this manner, the counter-rotating turbine600can be considered a seven stage turbine.

According to some embodiments there is a turbomachine characterized by both a high gear ratio and a high power gearbox. A high gear ratio gearbox means a gearbox with a gear ratio of above about 4:1. Examples of a high power gearbox include a gearbox adapted for transmitting power greater than 7 MW with output spool speed above, e.g., 1000 rpm, a gearbox adapted for transmitting power greater than 15 MW with output spool speed of about 1100 rpm, and a gearbox adapted for transmitting power greater than transmitting 22 MW with output spool speed of about 3500 rpm.

Each of the embodiments of turbomachines disclosed herein utilize a high gear ratio gearbox. Adoption of a gearbox having a high gear ratio presents unique challenges. One such challenge is determining the amount of oil that would need to circulate through the gearbox during operation, i.e., the high gear ratio gearbox's oil flow rate. The oil demand is significant when the engine requires a high gear ratio gearbox. Moreover, the estimated amount of oil flow for the high gear ratio gearbox is not well informed by, or capable of being estimated from, oil flow rates for an existing serviced engine. Starting from this basis, the inventors set out to calculate the oil flow demands for the different engine embodiments contemplated and disclosed herein, by consideration of the different features and performance characteristics, e.g., pitch line velocity and constants differentiating one gearbox configuration from another. The high gear ratio gearbox architectures considered include those described and disclosed herein (e.g.,FIGS. 9-13and the accompanying text, infra). These efforts accordingly involved factoring in specific characteristics of the gearboxes and the power transmission requirements for the gearbox to estimate the oil flow rates.

During the process of developing the aforementioned embodiments of turbomachines incorporating a high gear ratio gearbox, the inventors discovered, unexpectedly, that a good approximation of the high gear ratio gearbox oil flow rate may be made using only a relatively few engine parameters. This development is based on, among other things, the recognition that an oil flow rate through a gearbox is related to the expected power loss when transmitting power across a gearbox. From this initial recognition and other developments that were the by-product of studying several different engine configurations that included a power gearbox (including the configurations disclosed herein), the inventors ultimately discovered that a good approximation to the high gear ratio gearbox oil flow rate could be made based on a relationship among the turbomachine's gearbox gear ratio, net thrust, and fan diameter. The inventors refer to this relationship as a gearbox efficiency rating.

This discovery is quite beneficial. For example, with the gearbox efficiency rating having provided the engine oil flow requirements one can also estimate, for purposes of system integration, the type of oil-related secondary systems (e.g., sump, oil circuit, heat sinks, etc.) that would be included to support proper functioning of the selected high gear ratio gearbox; and/or to provide guidance on whether a particular engine architecture is beneficial or not, without requiring an entire team to complete the tedious and time-consuming process of developing a new gearbox from scratch. Therefore, the gearbox efficiency rating can improve the process of developing a turbomachine engine, which can ultimately result in improved turbomachines.

As indicated, the gearbox efficiency rating is a relationship based on a turbomachine's fan diameter (D), net thrust (T), and gear ratio of a high gear ratio gearbox. The gear efficiency rating, valid for gear ratios between about 4:1 and 14:1, may be expressed as Q(D1.56/T)1.56, where Q is measured at an inlet of the gearbox in gallons per minute at a max takeoff condition, D is measured in inches, and T is measured in pounds force at the max takeoff condition. In this manner, the gearbox efficiency rating defines a specific turbomachine engine configuration.

As used herein “net thrust” (T) equals the change of momementum of the bypass airflow plus the change of momentum of the core airflow and the burned fuel. Or stated another way, T=Wbyp(Vbyp−V0)+(Wcore+Wfuel) Vcore−WcoreV0, where Wbypis the mass flow rate of air of the bypass airflow, Vbypis the velocity of the bypass airflow, V0is the flight velocity, Wcoreis the mass flow rate of air of the core airflow, Wfuelis the mass flow rate of the burned fuel, and Vcoreis the velocity of the core airflow.

As indicated earlier, turbomachine engines, such as the turbofan engines100,200,300,400, comprise many variables and factors that affect their performance and/or operation. The interplay between the various components can make it particularly difficult to develop or select one component, especially when each of the components is at a different stage of completion. For example, one or more components may be nearly complete, yet one or more other components may be in an initial or preliminary phase where only one (or a few) parameters is known. Also, each component is subject to change often more than once over the development period, which can often last for many years (e.g., 5-15 years). These complex and intricate individual and collective development processes can be cumbersome and inefficient. For at least these reasons, there is a need for devices and methods that can provide a good estimate of, not only the basic configuration or sizing needed to achieve the desired performance benefits, but also to reflect the penalties or accommodations in other areas in order to realize the desired benefits.

According to another aspect of the disclosure, the gearbox efficiency rating may additionally provide a particularly useful indication of the efficiency and effectiveness of the engine during initial development, e.g., as a tool to accept or reject a particular configuration. Thus, the gearbox efficiency rating can be used, for example, to guide gearbox development. For example, the gearbox efficiency rating can be used to quickly and accurately determine the size of the gearbox that is suitable for a particular engine without requiring an individual or team to complete the tedious and time-consuming process of developing the gearbox from scratch. Therefore, the gearbox efficiency rating can also improve the process of developing a turbomachine engine.

FIGS. 7A-8illustrate exemplary ranges and/or values for gear efficiency rating.FIGS. 7A-7Cdisclose exemplary ranges of gear efficiency rating with respect to various gear ratios.FIG. 8discloses the gear ratio, oil flow, fan diameter, net thrust, and gearbox efficiency ratings for multiple exemplary turbomachine engines.

In some embodiments, the gearbox efficiency rating of a turbomachine engine is within a range of about 0.10-1.8 or 0.19-1.8 or 0.10-1.01. In certain embodiments, the gearbox efficiency rating is within a range of about 0.25-0.55 or about 0.29-0.51.FIG. 8provides the gear efficiency rating of several exemplary engines.

In some embodiments, the oil flow rate Q is within a range of about 5-55 gallons per minute. In certain embodiments, the oil flow rate Q is within a range of about 5.5-25 gallons per minute.FIG. 8also provides the oil flow rates of several exemplary engines.

As noted above, the oil flow rate Q is measured at an inlet of the gearbox in gallons per minute at a max takeoff condition. The inlet of the gearbox is the location at which the oil enters the gearbox from the oil supply line. As used herein “a max takeoff condition” means sea-level elevation, standard pressure, extreme hot day temperature, and a flight velocity of up to about 0.25 Mach.

As used herein, the term “extreme hot day temperature” means the extreme hot day temperature specified for a particular engine. This can include the extreme hot day temperature used for engine certification. Extreme hot day temperature can additionally or alternatively include temperatures of about 130-140° F.

In some embodiments, the fan diameter D is about 120-216 inches. In certain embodiments, the fan diameter D is about 120-192 inches.FIG. 8also provides the fan diameter of several exemplary engines.

In some embodiments, the net thrust T of the engine is within a range of about 10,000-100,000 pounds force. In particular embodiments, the net thrust T of the engine is within a range of about 12,000-30,000 pounds force.FIG. 8also provides the net thrust of several exemplary engines.

In some embodiments, the gearbox efficiency rating of a turbomachine engine can be configured in relation to the gear ratio (GR) of the gearbox. For example, in certain embodiments, a turbomachine engine can be configured such that the gearbox efficiency rating is greater than 0.015(GR1.4) and less than 0.034(GR1.5), as depicted inFIG. 7A. In other embodiments, a turbomachine engine can be configured such that the gearbox efficiency rating is greater than 0.02625(GR1.4) and less than 0.042(GR1.4).

For example,FIG. 8depicts several examplary engines with gearbox efficiency ratings that satisfy these relationships. Engine1is a turbomachine engine comprising a gearbox with a gear ratio of 10.5:1 and a gearbox efficiency rating within a range of 0.40-1.16, specifically 1.02. Engine2, Engine3, and Engine4are turbomachine engines comprising gearboxes with a gear ratio of 7:1 and the gearbox efficiency ratings within a range of 0.23-0.63, that is 0.51, 0.42, and 0.41, respectively. Engine5is a turbomachine engine comprising a gearbox with a gear ratio of 5.1:1 and a gearbox efficiency rating within a range of 0.15-0.39, specifically 0.29. Engine6is a turbomachine engine comprising a gearbox with a gear ratio of 4.1:1 and a gearbox efficiency rating within a range of 0.11-0.28, specifically 0.21. Engines7-19provide additional examples with specific gearbox efficiency ratings. Ranges for the gearbox efficiency ratings of Engines7-19can be determined using the equations above and/or the charts ofFIGS. 7A-7C.

As another example, a turbomachine engine comprising a gearbox with a gear ratio of 4.5:1 can be configured such that the gearbox efficiency rating is within a range of 0.12-0.32. As another example, a turbomachine engine comprising a gearbox with a gear ratio of 6:1 can be configured such that the gearbox efficiency rating within a range of 0.18-0.50. As another example, a turbomachine engine comprising a gearbox with a gear ratio of 9:1 can be configured such that the gearbox efficiency rating within a range of 0.33-0.92. As another example, a turbomachine engine comprising a gearbox with a gear ratio of 11:1 can be configured such that the gearbox efficiency rating within a range of 0.43-1.24. As another example, a turbomachine engine comprising a gearbox with a gear ratio of 12:1 can be configured such that the gearbox efficiency rating within a range of 0.49-1.41. As yet another example, a turbomachine engine comprising a gearbox with a gear ratio of 14:1 can be configured such that the gearbox efficiency rating is within a range of 0.60-1.78.

In some instances, a turbomachine engine can comprise a gearbox with a gear ratio of 5-6, 7-8, 9-10, 11-12, or 13-14. In other instances, a turbomachine engine can comprise a gearbox with a gear ratio of 5-7, 8-10, 11-13. In yet other embodiments, a turbomachine engine can comprise a gearbox with a gear ratio of 7-10 or 11-14. Below is a table with several exemplary gearbox efficiency ratings with respect to several exemplary gear ratios.

In some embodiments, a turbomachine engine can be configured such that the gearbox efficiency rating is greater than 0.023 (GR1.5) and less than 0.034(GR1.5), as depicted inFIG. 7B. In particular instances, the gearbox efficiency rating can be about 0.0275(GR1.5). These embodiments can be particularly advantageous, for example, with engines comprising an epicyclic gearbox (e.g., star and/or planet configuration).

In other embodiments, a turbomachine engine can be configured such that the gearbox efficiency rating is greater than 0.015(GR1.4) and less than 0.025(GR1.4), as depicted inFIG. 7C. In particular instances, the gearbox efficiency rating can be about 0.02 (GR1.4). These embodiments can be particularly advantageous, for example, with engines comprising a non-epicyclic gearbox (e.g., compound gearboxes).

It should be noted gearbox efficiency rating values disclosed herein are approximate values. Accordingly, the disclosed gearbox efficiency rating values include values within five percent of the listed values.

As noted above, the gearbox efficiency rating can define a specific engine configuration and/or can be used when developing a gearbox for a turbomachine engine. For example, in some instances, the gearbox efficiency rating can be used to determine the size and/or oil flow rate of a gearbox. Assuming that a desired gear ratio of the gearbox is known, along with the fan diameter, and the net thrust of the engine, the gearbox efficiency ratings depicted in the charts ofFIG. 7A-7Ccan be used to determine an acceptable oil flow rate. In some embodiments, the equation below can be used to determine an acceptable range of oil flow rates (Q) for the gearbox. The determined oil flow rate Q can be used, for example, to aid in the configuration of the gearbox. In some instances, one or more other parameters (e.g., the gearbox efficiency rating) can also aid in the configuration of the gearbox.

For example, a gearbox for a turbomachine engine can be configured using the following exemplary method. With reference toFIG. 8, Engine1comprises an unducted fan and can be configured similar to the engine200. Engine1comprises a fan diameter of 188.6 inches and a net thrust of 25,503 pounds force at a max takeoff condition. Engine1further comprises a five stage low-pressure turbine. The desired gear ratio for the gearbox of Engine1is about 10.5:1. Based on this information, oil flow rate Q of the gearbox of Engine1should be about 8-24 gallons per minute at a max takeoff condition.

FIG. 9schematically depicts a gearbox700that can be used, for example, with Engine1. The gearbox700comprises a two-stage star configuration.

The first stage of the gearbox700includes a first-stage sun gear702, a first-stage carrier704housing a plurality of first-stage star gears, and a first-stage ring gear706. The first-stage sun gear702can be coupled to a low-speed shaft708, which in turn is coupled to the low-pressure turbine of Engine1. The first-stage sun gear702can mesh with the first-stage star gears, which mesh with the first-stage ring gear. The first-stage carrier704can be fixed from rotation by a support member710.

The second stage of the gearbox700includes a second-stage sun gear712, a second-stage carrier714housing a plurality of second-stage star gears, and a second-stage ring gear716. The second-stage sun gear712can be coupled to a shaft718which in turn is coupled to the first-stage ring gear706. The second-stage carrier714can be fixed from rotation by a support member720. The second-stage ring gear716can be coupled to a fan shaft722.

In some embodiments, each stage of the gearbox700can comprise five star gears. In other embodiments, the gearbox700can comprise fewer or more than five star gears in each stage. In some embodiments, the first-stage carrier can comprise a different number of star gears than the second-stage carrier. For example, the first-carrier can comprise five star gears, and the second-stage carrier can comprise three star gears, or vice versa.

Based on the configuration of the gearbox700and the calculated oil flow rate of 8-24 gallons per minute, which is based on the gearbox efficiency rating, the gearbox700can comprise a radius R1. The size of the gearbox, including the radius R1, can be configured such that the oil flow rate at the inlet of the gearbox700at a max takeoff condition is about 8-24 gallons per minute or about 16-24 gallons per minute (e.g., 20.9 gpm). In some embodiments, the radius R1of the gearbox700can be about 16-19 inches. In other embodiments, the radius R1of the gearbox700can be about 22-24 inches. In other embodiments, the radius R1of the gearbox700can be smaller than 16 inches or larger than 24 inches.

As another example, Engine2(FIG. 8) comprises an unducted fan and can be configured similar to the engine200. Engine3comprises a fan diameter of 188.6 inches and a net thrust of 25,000 pounds force at a max takeoff condition. Engine2further comprises a 3-7 stage low-pressure turbine. The desired gear ratio for the gearbox of Engine2is about 7:1. Based on this information, oil flow rate Q of the gearbox of Engine2should be about 4-13 gallons per minute or about 8-13 gallons per minute (e.g., 10.06) at a max takeoff condition.

FIG. 10schematically depicts a gearbox800that can be used, for example, with Engine2. The gearbox800comprises a single-stage star configuration. The gearbox800includes a sun gear802, a carrier804housing a plurality of star gears (e.g., 3-5 star gears), and a ring gear806. The sun gear802can mesh with the star gears, and the star gears can mesh with the ring gear806. The sun gear802can be coupled to a low-speed shaft808, which in turn is coupled to the low-pressure turbine of Engine2. The carrier804can be fixed from rotation by a support member810. The ring gear806can be coupled to a fan shaft812.

Based on the configuration of the gearbox800and the calculated oil flow rate of 4-13 gallons per minute, which is based on the gearbox efficiency rating, the gearbox800can comprise a radius R2. The size of the gearbox, including the radius R2, can be configured such that the oil flow rate at the inlet of the gearbox800at a max takeoff condition is 7-13 gallons per minute (e.g., 10.1 gpm). In some embodiments, the radius R2of the gearbox800can be about 18-23 inches. In other embodiments, the radius R2of the gearbox700can be smaller than 18 inches or larger than 23 inches.

As another example, Engine3(FIG. 8) comprises an unducted fan and can be configured similar to the engine200. Engine3comprises a fan diameter of 142.8 inches and a net thrust of 12,500 pounds force at a max takeoff condition. Engine3further comprises a 3-7 stage low-pressure turbine. The desired gear ratio for the gearbox of Engine3is about 7:1. Based on this information, oil flow rate Q of the gearbox of Engine3should be about 3-9 gallons per minute or about 5-9 gallons per minute (e.g., 6 gpm) at a max takeoff condition.

FIG. 11schematically depicts a gearbox900that can be used, for example, with Engine3. The gearbox800comprises a single-stage star configuration. The gearbox900includes a sun gear902, a carrier904housing a plurality of star gears (e.g., 3-5 star gears), and a ring gear906. The sun gear902can mesh with the star gears, and the star gears can mesh with the ring gear906. The sun gear902can be coupled to a low-speed shaft908, which in turn is coupled to the low-pressure turbine of Engine3. The carrier904can be fixed from rotation by a support member910. The ring gear906can be coupled to a fan shaft912.

Based on the configuration of the gearbox900and the calculated oil flow rate of 5-9 gallons per minute, which is based on the gearbox efficiency rating, the gearbox900can comprise a radius R3. The size of the gearbox, including the radius R3, can be configured such that the oil flow rate at the inlet of the gearbox900at a max takeoff condition is 3-9 gallons per minute (e.g., 6 gpm). In some embodiments, the radius R3of the gearbox900can be about 10-13 inches. In other embodiments, the radius R3of the gearbox900can be smaller than 10 inches or larger than 13 inches.

Engine4comprises an unducted fan and can be configured similar to the engine200. Engine4comprises a fan diameter of 188.4 inches and a net thrust of 25,000 pounds force at a max takeoff condition. Engine4further comprises a counter-rotating low-pressure turbine (e.g., similar to the counter-rotating turbine500or the counter-rotating turbine600). The desired gear ratio for the gearbox of Engine4is about 7:1. Based on this information, oil flow rate Q of the gearbox of Engine4should be about 4-13 gallons per minute or about 7-13 gallons per minute (e.g., 8.1 gpm) at a max takeoff condition.

FIG. 12schematically depicts a gearbox1000that can be used, for example, with Engine4. The gearbox1000comprises a two-stage configuration in which the first stage is a star configuration and the second stage is a planet configuration.

The first stage of the gearbox1000includes a first-stage sun gear1002, a first-stage star carrier1004comprising a plurality of first-stage star gears (e.g., 3-5 star gears), and a first-stage ring gear1006. The first-stage sun gear1002can mesh with the first-stage star gears, and the first-stage star gears can mesh with the first-stage ring gear1006. The first-stage sun gear1002can be coupled to a higher-speed shaft1008of the low spool, which in turn is coupled to the inner blades of the low-pressure turbine of Engine4. The first-stage star carrier1004can be fixed from rotation by a support member1010.

The second stage of the gearbox1000includes a second-stage sun gear1012, a second-stage planet carrier1014comprising a plurality of second-stage planet gears (e.g., 3-5 planet gears), and a second-stage ring gear1016. The second-stage sun gear1012can mesh with the second-stage planet gears. The second-stage planet carrier1014can be coupled to the first-stage ring gear1006. The second-stage sun gear1012can be coupled to a lower-speed shaft1018of the low spool, which in turn is coupled to the outer blades of the low-pressure turbine of Engine4. The second-stage planet carrier1014can be coupled to the first-stage ring gear1006. The second-stage planet carrier1014can also be coupled to a fan shaft1020. The second-stage ring gear1016can be fixed from rotation by a support member1022.

In some embodiments, each stage of the gearbox1000can comprise three star/planet gears. In other embodiments, the gearbox1000can comprise fewer or more than three star/planet gears in each stage. In some embodiments, the first-stage carrier can comprise a different number of star gears than the second-stage carrier has planet gears. For example, the first-carrier can comprise five star gears, and the second-stage carrier can comprise three planet gears, or vice versa.

Since the first stage of the gearbox1000is coupled to the higher-speed shaft1008of the low spool and the second stage of the gearbox1000is coupled to the lower-speed shaft1018of the low spool, the gear ratio of the first stage of the gearbox1000can 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 comprise a gear ratio of 4.1-14, and the second stage of the gearbox can comprise a gear ratio that is less than the gear ratio of the first stage of the gearbox. In particular embodiments, the first stage of the gearbox can comprise a gear ratio of 7, and the second stage of the gearbox can comprise a gear ratio of 6.

In some embodiments, an engine comprising the gearbox1000can be configured such that the higher-speed shaft1008provides about 50% of the power to the gearbox1000and the lower-speed shaft1018provides about 50% of the power to the gearbox1000. In other embodiments, an engine comprising the gearbox1000can be configured such that the higher-speed shaft1008provides about 60% of the power to the gearbox1000and the lower-speed shaft1018provides about 40% of the power to the gearbox1000.

Based on the configuration of the gearbox1000and the calculated oil flow rate of 4-13 gallons per minute, which is based on the gearbox efficiency rating, the gearbox1000can comprise a radius R4. The size of the gearbox, including the radius R4, can be configured such that the oil flow rate at the inlet of the gearbox1000at a max takeoff condition is 7-13 gallons per minute (e.g., 8.1 gpm). In some embodiments, the radius R4of the gearbox1000can be about 18-22 inches. In other embodiments, the radius R4of the gearbox700can be smaller than 18 inches or larger than 22 inches.

Thus, as illustrated by the examples disclosed herein, a gearbox efficiency rating can characterize or define a specific engine and/or gearbox configuration. As such, turbomachine engines can be quickly and accurately configured by utilizing the gearbox efficiency rating and/or its related parameters. In this manner, the gearbox efficiency rating disclosed herein provides one or more significant advantages over known turbomachine engines and/or known methods of developing turbomachine engines.

FIG. 13depicts a gearbox1100that can be used, for example, with the engines disclosed herein (e.g., the engines100,200,400). The gearbox1100is configured as a compound star gearbox. The gearbox1100comprises a sun gear1102and a star carrier1104, which includes a plurality of compound star gears having one or more first portions1106and one or more second portions1108. The gearbox1100further comprises a ring gear1110. The sun gear1102can also mesh with the first portions1106of the star gears. The star carrier can be fixed from rotation via a support member1114. The second portions1108of the star gears can mesh with the ring gear1110. The sun gear1102can be coupled to a low-pressure turbine via the turbine shaft1112. The ring gear1110can be coupled to a fan shaft1116.

The gear assemblies shown and described herein can be used with any suitable engine. For example, althoughFIG. 4shows an optional ducted fan and optional fan duct (similar to that shown inFIG. 2), it should be understood that such gear assemblies can be used with other ducted turbofan engines (e.g., the engine300) and/or other open rotor engines that do not have one or more of such structures.

Embodiments of the gear assemblies depicted and described herein may provide for gear ratios and arrangements that fit within the L/Dcoreconstraints of the disclosed engines. In certain embodiments, the gear assemblies depicted and described in regard toFIGS. 9-13allow for gear ratios and arrangements providing for rotational speed of the fan assembly corresponding to one or more ranges of cruise altitude and/or cruise speed provided above.

Various embodiments of the gear assembly provided herein may allow for gear ratios of up to 14:1. Still various embodiments of the gear assemblies provided herein may allow for gear ratios of at least 4.1:1 or 4.5:1. Still yet various embodiments of the gear assemblies provided herein allow for gear ratios of 6:1 to 12:1.FIG. 8also provides the gear ratio of several exemplary engines. It should be appreciated that embodiments of the gear assemblies provided herein may allow for large gear ratios and within constraints such as, but not limited to, length of the engine, maximum diameter (Dcore) of the engine100, cruise altitude of up to 65,000 ft, and/or operating cruise speed of up to Mach 0.85, or combinations thereof.

Various exemplary gear assemblies are shown and described herein. These gear assemblies may be utilized with any of the exemplary engines and/or any other suitable engine for which such gear assemblies may be desirable. In such a manner, it will be appreciated that the gear assemblies disclosed herein may generally be operable with an engine having a rotating element with a plurality of rotor blades and a turbomachine having a turbine and a shaft rotatable with the turbine. With such an engine, the rotating element (e.g., fan assembly104) may be driven by the shaft (e.g., low-speed shaft146) of the turbomachine through the gear assembly.

Although the exemplary gear assemblies shown are mounted at a forward location (e.g., forward from the combustor and/or the low-pressure compressor), in other embodiments, the gear assemblies described herein can be mounted at a aft location (e.g., aft of the combustor and/or the low-pressure turbine).

Portions of a lubricant system1200are depicted schematically inFIG. 14. The lubrication system1200can be a component of the turbomachine engines disclosed herein and/or can be coupled to the various gearboxes disclosed herein. For example,FIG. 1schematically illustrates the lubricant system coupled to the turbomachine engine100and the gearbox102. A series of lubricant conduits1203can interconnect multiple elements of the lubricant system1200and/or engine components, thereby providing for provision or circulation of the lubricant throughout the lubricant system and any engine components coupled thereto (e.g., a gearbox, bearing compartments, etc.).

It should be understood that the organization of the lubricant system1200as shown is by way of example only to illustrate an exemplary system for a turbomachine engine for circulating lubricant for purposes such as lubrication or heat transfer. Any organization for the lubricant system1200is contemplated, with or without the elements as shown, and/or including additional elements interconnected by any necessary conduit system.

Referring again toFIG. 14, the lubricant system1200includes a lubricant reservoir1202configured to store a coolant or lubricant, including organic or mineral oils, synthetic oils, or fuel, or mixtures or combinations thereof. A supply line1204and a scavenge line1206are fluidly coupled to the reservoir1202and collectively form a lubricant circuit to which the reservoir1202and component1210(e.g., a gearbox) can be fluidly coupled. The component1210can be supplied with lubrication by way of a fluid coupling with the supply line1204and can return the supplied lubricant to the reservoir1202by fluidly coupling to the scavenge line1206. More specifically, a component supply line1211can be fluidly coupled between the supply line1204and the component1210. It is further contemplated that multiple types of lubricant can be provided in other lines not explicitly shown, but are nonetheless included in the lubricant system1200.

Optionally, at least one heat exchanger1205can be included in the lubricant system1200. The heat exchanger1205can include a fuel/lubricant (fuel-to-lubricant) heat exchanger, an oil/lubricant heat exchanger, an air cooled oil cooler, and/or other means for exchanging heat. For example, a fuel/lubricant heat exchanger can be used to heat or cool engine fuel with lubricant passing through the heat exchanger. In another example, a lubricant/oil heat exchanger can be used to heat or cool additional lubricants passing within the turbomachine engine, fluidly separate from the lubricant passing along the lubricant system1200. Such a lubricant/oil heat exchanger can also include a servo/lubricant heat exchanger. Optionally, a second heat exchanger (not shown) can be provided along the exterior of the core engine, downstream of the outlet guide vane assembly. The second heat exchanger can be an air/lubricant heat exchanger, for example, adapted to convectively cool lubricant in the lubricant system1200utilizing the airflow passing through an outlet guide vane assembly of the turbomachine engine.

A pump1208can be provided in the lubricant system1200to aid in recirculating lubricant from the reservoir1202to the component1210via the supply line1204. For example, the pump1208can be driven by a rotating component of the turbine engine10, such as a high-pressure shaft or a low-pressure shaft of a turbomachine engine.

Lubricant can be recovered from the component1210by way of the scavenge line1206and returned to the reservoir1202. In the illustrated example, the pump1208is illustrated along the supply line1204downstream of the reservoir1202. The pump1208can be located in any suitable position within the lubricant system1200, including along the scavenge line1206upstream of the reservoir1202. In addition, while not shown, multiple pumps can be provided in the lubricant system1200.

In some embodiments, a bypass line1212can be fluidly coupled to the supply line1204and scavenge line1206in a manner that bypasses the component1210. In such embodiments, a bypass valve1215is fluidly coupled to the supply line1204, component supply line1211, and bypass line1212. The bypass valve1215is configured to control a flow of lubricant through at least one of the component supply line1211or the bypass line1212. The bypass valve1215can include any suitable valve including, but not limited to, a differential thermal valve, rotary valve, flow control valve, and/or pressure safety valve. In some embodiments, a plurality of bypass valves can be provided.

During operation, a supply flow1220can move from the reservoir1202, through the supply line1204, and to the bypass valve1215. A component input flow1222can move from the bypass valve1215through the component supply line1211to an inlet of the component1210. A scavenge flow1224can move lubricant from an outlet of the component1210through the scavenge line1206and back to the reservoir1202. Optionally, a bypass flow1226can move from the bypass valve1215through the bypass line1212and to the scavenge line1206. The bypass flow1226can mix with the scavenge flow1224and define a return flow1228moving toward the lubricant reservoir1202.

In one example where no bypass flow exists, it is contemplated that the supply flow1220can be the same as the component input flow1222and that the scavenge flow1224can be the same as the return flow1228. In another example where the bypass flow1226has a nonzero flow rate, the supply flow1220can be divided at the bypass valve1215into the component input flow1222and bypass flow1226. It will also be understood that additional components, valves, sensors, or conduit lines can be provided in the lubricant system1200, and that the example shown inFIG. 14is simplified with a single component1210for purposes of illustration.

The lubricant system1200can further include at least one sensing position at which at least one lubricant parameter can be sensed or detected. The at least one lubricant parameter can include, but is not limited to, a flow rate, a temperature, a pressure, a viscosity, a chemical composition of the lubricant, or the like. In the illustrated example, a first sensing position1216is located in the supply line1204upstream of the component1210, and a second sensing position1218is located in the scavenge line1206downstream of the component1210.

In one example, the bypass valve1215can be in the form of a differential thermal valve configured to sense or detect at least one lubricant parameter in the form of a temperature of the lubricant. In such a case, the fluid coupling of the bypass valve1215to the first and second sensing positions1216,1218can provide for bypass valve1215sensing or detecting the lubricant temperature at the sensing positions1216,18as lubricant flows to or from the bypass valve1215. The bypass valve1215can be configured to control the component input flow1222or the bypass flow1226based on the sensed or detected temperature.

It is contemplated that the bypass valve1215, supply line1204, and bypass line1212can at least partially define a closed loop control system for the component1210. As used herein, a “closed loop control system” will refer to a system having mechanical or electronic components that can automatically regulate, adjust, modify, or control a system variable without manual input or other human interaction. Such closed loop control systems can include sensing components to sense or detect parameters related to the desired variable to be controlled, and the sensed or detected parameters can be utilized as feedback in a “closed loop” manner to change the system variable and alter the sensed or detected parameters back toward a target state. In the example of the lubricant system1200, the bypass valve1215(e.g. mechanical or electrical component) can sense a parameter, such as the lubricant parameter (e.g. temperature), and automatically adjust a system variable, e.g., flow rate to either or both of the bypass line1212or component1210, without need of additional or manual input. In one example, the bypass valve can be automatically adjustable or self-adjustable such as a thermal differential bypass valve. In another example, the bypass valve can be operated or actuated via a separate controller. It will be understood that a closed loop control system as described herein can incorporate such a self-adjustable bypass valve or a controllable bypass valve.

Turning toFIG. 15, a portion of the lubricant system1200is illustrated supplying lubricant to a particular component1210in the form of a gearbox1250within a turbomachine engine. The gearbox can be any of the gearboxes disclosed herein. The gearbox1250can include an input shaft1252, an output shaft1254, and a gear assembly1255. In one example, the gear assembly1255can be in the form of an epicyclic gear assembly as known in the art having a ring gear, sun gear, and at least one planet/star gear. An outer housing1256can at least partially surround the gear assembly1255and form a structural support for the gears and bearings therein. Either or both of the input and output shafts1252,1254can be coupled to the turbomachine engine. In one example, the input and output shafts1252,1254can be utilized to decouple the speed of the low-pressure turbine from the low-pressure compressor and/or the fan, which can, for example, improve engine efficiency.

The supply line1204can be fluidly coupled to the gearbox1250, such as to the gear assembly1255, to supply lubricant to gears or bearings to the gearbox1250during operation. The scavenge line1206can be fluidly coupled to the gearbox1250, such as to the gear assembly1255or outer housing1256, to collect lubricant. The bypass line1212can be fluidly coupled to the bypass valve1215, supply line1204, and scavenge line1206as shown. A return line1214can also be fluidly coupled to the bypass valve1215, such as for directing the return flow1228to the lubricant reservoir1202for recirculation. While not shown inFIG. 15for brevity, the lubricant reservoir1202, the heat exchanger1205, and/or the pump1208(FIG. 14) can also be fluidly coupled to the gearbox1250. In this manner, the supply line1204, bypass line1212, scavenge line1206, and return line1214can at least partially define a recirculation line1230for the lubricant system1200.

The supply flow1220divides at the bypass line into the component input flow1222and the bypass flow1226. In the example shown, the bypass valve1215is in the form of a differential thermal valve that is fluidly coupled to the first and second sensing positions1216,1218.

Lubricant flowing proximate the first and second sensing positions1216,1218provides the respective first and second outputs1241,1242indicative of the temperature of the lubricant at those sensing positions1216,1218. It will be understood that the supply line1204is thermally coupled to the bypass line1212and bypass valve1215such that the temperature of the fluid in the supply line1204proximate the first sensing position1216is approximately the same as fluid in the bypass line1212adjacent the bypass valve1215. Two values being “approximately the same” as used herein will refer to the two values not differing by more than a predetermined amount, such as by more than 20%, or by more than 5 degrees, in some examples. In this manner, the bypass valve1215can sense the lubricant temperature in the supply line1204and scavenge line1206via the first and second outputs1241,1242. It can be appreciated that the bypass line1212can form a sensing line for the valve1215to sense the lubricant parameter, such as temperature, at the first sensing position1216.

During operation of the turbomachine engine, the lubricant temperature can increase within the gearbox1250, such as due to heat generation of the gearbox1250, and throughout the lubricant system1200. In one example, if a lubricant temperature exceeds a predetermined threshold temperature at either sensing position1216,1218, the bypass valve1215can automatically increase the component input flow1222, e.g. from the supply line1204to the gearbox1250, by decreasing the bypass flow1226. Such a predetermined threshold temperature can be any suitable operating temperature for the gearbox1250, such as about 300° F. in some examples. Increasing the component input flow1222can provide for cooling of the gearbox1250, thereby reducing the lubricant temperature sensed in the various lines1204,1206,1212,1214as lubricant recirculates through the lubricant system1200.

In another example, if a temperature difference between the sensing positions1216,1218exceeds a predetermined threshold temperature difference, the bypass valve can automatically increase the component input flow1222by decreasing the bypass flow1226. Such a predetermined threshold temperature difference can be any suitable operating temperature for the gearbox1250, such as about 70° F., or differing by more than 30%, in some examples. In yet another example, if a temperature difference between the sensing positions1216,1218is below the predetermined threshold temperature difference, the bypass valve can automatically decrease the component input flow1222or increase the bypass flow1226. In this manner the lubricant system1200can provide for the gearbox to operate with a constant temperature difference between the supply and scavenge lines1204,1206.

1. A turbomachine engine comprising a fan assembly including a plurality of fan blades, a vane assembly including a plurality of vanes, a core engine including one or more compressor sections and one or more turbine sections, a gearbox including an input and an output, and a gearbox efficiency rating of 0.10-1.8. The input is coupled to the one or more turbine sections of the core engine and comprises a first rotational speed, the output is coupled to the fan assembly and has a second rotational speed, and a gear ratio of the first rotational speed to the second rotational speed is within a range of 4.1-14.0. The gearbox efficiency rating equals Q(D{circumflex over ( )}1.56/T){circumflex over ( )}1.53, where Q is a gearbox oil flow rate at an inlet of the gearbox measured in gallons per minute at a max takeoff condition, D is a diameter of the fan blades measured in inches, and T is a net thrust of the turbomachine engine measured in pounds force at the max takeoff condition.

2. The turbomachine engine of any clause herein, wherein the gearbox efficiency rating is 0.10-1.01.

3. The turbomachine engine of any clause herein, wherein the gearbox efficiency rating is 0.19-1.8

4. The turbomachine engine of any clause herein, wherein the gear ratio is within a range of 4.5-12.0.

5. The turbomachine engine of any clause herein, wherein the gear ratio is within a range of 6.0-11.0.

6. The turbomachine engine of any clause herein, wherein Q is within a range of 5-55 gallons per minute.

7. The turbomachine engine of any clause herein, wherein Q is within a range of 6-36 gallons per minute.

8. The turbomachine engine of any clause herein, wherein D is 120-216 inches.

9. The turbomachine engine of any clause herein, wherein D is 120-192 inches.

10. The turbomachine engine of any clause herein, wherein T is within a range of 10,000-100,000 pounds force.

11. The turbomachine engine of any clause herein, wherein T is within a range of 12,000-30,000 pounds force.

12. The turbomachine engine of any clause herein, wherein the gearbox is an epicyclic gearbox comprising a sun gear, a plurality of planet gears, and a ring gear, wherein the sun gear is the input, and wherein the ring gear is the output.

13. The turbomachine engine of any clause herein, wherein the gearbox is an epicyclic gearbox comprising a sun gear, a plurality of planet gears, and a ring gear, wherein the sun gear is the input, wherein the planet gears are coupled to a planet carrier, and wherein the planet carrier is the output.

14. The turbomachine engine of any clause herein, wherein the gearbox is a multi-stage gearbox.

15. The turbomachine engine of any clause herein, wherein the gearbox is a two-stage gearbox.

16. The turbomachine engine of any clause herein, wherein the gearbox is a compound gearbox.

17. A turbomachine engine comprising a fan assembly including a plurality of fan blades, a vane assembly including a plurality of vanes, a core engine including a low-pressure compressor, a high-pressure compressor, a combustor, a high-pressure turbine, and a low-pressure turbine, a gearbox including an input and an output, and a gearbox efficiency rating of 0.12-1.8. The input is coupled to the low-pressure turbine and comprises a first rotational speed, the output is coupled to the fan assembly and has a second rotational speed, and a gear ratio of the first rotational speed to the second rotational speed is within a range of 4.5-14.0. The gearbox efficiency rating equals Q(D{circumflex over ( )}1.56/T){circumflex over ( )}1.53, where Q is a gearbox oil flow rate at an inlet of the gearbox measured in gallons per minute at a max takeoff condition, D is a diameter of the fan blades measured in inches, and T is a net thrust of the turbomachine engine measured in pounds force at the max takeoff condition.

18. The turbomachine engine of any clause herein, wherein Q is within a range of 5-55 gallons per minute.

19. The turbomachine engine of any clause herein, wherein D is 120-216 inches.

20. The turbomachine engine of any clause herein, wherein T is within a range of 10,000-100,000 pounds force.

21. The turbomachine engine of any clause herein, wherein the gearbox is an epicyclic gearbox comprising a star gear configuration.

22. The turbomachine engine of any clause herein, wherein the gearbox is an epicyclic gearbox comprising a planet gear configuration.

23. The turbomachine engine of any clause herein, wherein the fan assembly comprises 8-20 fan blades.

24. The turbomachine engine of any clause herein, wherein the low-pressure compressor comprises 1-8 stages.

25. The turbomachine engine of any clause herein, wherein the high-pressure compressor comprises 8-15 stages.

26. The turbomachine engine of any clause herein, wherein the high-pressure turbine comprises 1-2 stages.

27. The turbomachine engine of any clause herein, wherein the low-pressure turbine comprises 3-7 stages.

28. The turbomachine engine of any clause herein, wherein the low-pressure turbine is a counter-rotating low-pressure turbine comprising inner blade stages and outer blade stages, wherein the inner blade stages extend radially outwardly from an inner shaft, and wherein the outer blade stages extend radially inwardly from an outer drum.

29. The turbomachine engine of any clause herein, wherein the counter-rotating low-pressure turbine comprises four inner blade stages and three outer blade stages.

30. The turbomachine engine of any clause herein, wherein the counter-rotating low-pressure turbine comprises three inner blade stages and three outer blade stages.

31. A turbomachine engine comprising a fan assembly including a plurality of fan blades, a vane assembly including a plurality of vanes, a core engine including a low-pressure compressor, a high-pressure compressor, a combustor, a high-pressure turbine, and a low-pressure turbine. The turbomachine engine also includes a gearbox having a gear ratio within a range of 6.0-12.0, and a gearbox efficiency rating of 0.18-1.41.

32. The turbomachine engine of any clause herein, wherein a gearbox oil flow rate at an inlet of the gearbox is within a range of 5-55 gallons per minute at a max takeoff condition.

33. The turbomachine engine of any clause herein, wherein a diameter of the fan blades is 120-216 inches.

34. The turbomachine engine of any clause herein, wherein a net thrust of the turbomachine engine is within a range of 10,000-100,000 pounds force at a max takeoff condition.

35. A turbomachine engine comprises an unducted fan assembly, a core engine, a vane assembly, a gearbox, and a gearbox efficiency rating. The unducted fan assembly includes a single row of fan blades. The core engine including one or more compressor sections and one or more turbine sections. The vane assembly includes a single row of vanes. The vanes are disposed aft of the fan blades and comprise fixed end portions and free end portions. The fixed end portions are coupled to the core engine, and the free end portions are spaced radially outwardly from the core engine. The gearbox includes an input and an output. The input is coupled to the one or more turbine sections of the core engine and comprises a first rotational speed, the output is coupled to the unducted fan assembly and has a second rotational speed, and a gear ratio of the first rotational speed to the second rotational speed is within a range of 4.1-14.0. The gearbox efficiency rating is 0.10-1.8. The gearbox efficiency rating equals

Q⁡(D1.56T)1.53,
where Q is a gearbox oil flow rate at an inlet of the gearbox measured in gallons per minute at a max takeoff condition, D is a diameter of the fan blades measured in inches, and T is a net thrust of the turbomachine engine measured in pounds force at the max takeoff condition.

36. The turbomachine engine of any clause herein, further comprising a pitch change mechanism coupled to the unducted fan assembly.

37. The turbomachine engine of any clause herein, further comprising a pitch change mechanism coupled to the vane assembly.

38. The turbomachine engine of any clause herein, wherein the gearbox efficiency rating is 0.25-1.15.

39. The turbomachine engine of any clause herein, wherein the gear ratio is within a range of 4.5-12.0.

40. The turbomachine engine of any clause herein, wherein the gear ratio is within a range of 6.0-11.0.

41. The turbomachine engine of any clause herein, wherein Q is within a range of 5-55 gallons per minute.

42. The turbomachine engine of any clause herein, wherein D is 120-216 inches.

43. The turbomachine engine of any clause herein, wherein T is within a range of 10,000-100,000 pounds force.

44. The turbomachine engine of any clause herein, wherein the gearbox is an epicyclic gearbox comprising a sun gear, a plurality of planet gears, and a ring gear, wherein the sun gear is the input, and wherein the ring gear is the output.

45. The turbomachine engine of any clause herein, wherein the gearbox is an epicyclic gearbox comprising a sun gear, a plurality of planet gears, and a ring gear, wherein the sun gear is the input, wherein the planet gears are coupled to a planet carrier, and wherein the planet carrier is the output.

46. The turbomachine engine of any clause herein, wherein the gearbox is a multi-stage gearbox.

47. The turbomachine engine of any clause herein, wherein the gearbox is a compound gearbox.

48. A turbomachine engine comprises an unducted fan assembly, an unducted vane assembly, a ducted fan assembly, a core engine, a gearbox, and a gearbox efficiency rating. The unducted fan assembly includes a plurality of first fan blades. The unducted vane assembly including a plurality of vanes, and the vanes are positioned aft of the first fan blades. The ducted fan assembly includes a plurality of second fan blades, and the ducted fan assembly is positioned aft of the unducted fan assembly and radially inwardly from the unducted vane assembly. The core engine including a low-pressure compressor, a high-pressure compressor, a combustor, a high-pressure turbine, and a low-pressure turbine. The gearbox is coupled to the low-pressure turbine and the unducted fan assembly. The gearbox comprises a gear ratio of 4.1-14.0 and is configured such that the unducted fan assembly rotates slower than low-pressure turbine. The gearbox efficiency rating is 0.10-1.8. The gearbox efficiency rating equals

Q⁡(D1.56T)1.53,
where Q is a gearbox oil flow rate at an inlet of the gearbox measured in gallons per minute at a max takeoff condition, D is a diameter of the fan blades measured in inches, and T is a net thrust of the turbomachine engine measured in pounds force at the max takeoff condition.

49. The turbomachine engine of any clause herein, wherein the gearbox efficiency rating is 0.10-1.01.

50. The turbomachine engine of any clause herein, wherein the gearbox efficiency rating is 0.19-1.8.

51. The turbomachine engine of any clause herein, wherein Q is within a range of 5-40 gallons per minute.

52. The turbomachine engine of any clause herein, wherein D is 140-192 inches.

53. The turbomachine engine of any clause herein, wherein T is within a range of 10,000-40,000 pounds force.

54. The turbomachine engine of any clause herein, wherein the gearbox is an epicyclic gearbox comprising a star gear configuration.

55. The turbomachine engine of any clause herein, wherein the gearbox is an epicyclic gearbox comprising a planet gear configuration.

56. The turbomachine engine of any clause herein, wherein the unducted fan assembly comprises 8-14 fan blades.

57. The turbomachine engine of any clause herein, wherein the low-pressure compressor comprises 1-2 stages.

58. The turbomachine engine of any clause herein, wherein the high-pressure compressor comprises 10-11 stages.

59. The turbomachine engine of any clause herein, wherein the high-pressure turbine comprises two stages.

60. The turbomachine engine of any clause herein, wherein the low-pressure turbine comprises 6-7 stages.

61. The turbomachine engine of any clause herein, wherein the low-pressure turbine is a counter-rotating low-pressure turbine comprising inner blade stages and outer blade stages, wherein the inner blade stages extend radially outwardly from an inner shaft, and wherein the outer blade stages extend radially inwardly from an outer drum.

62. The turbofan engine of any clause herein, wherein the unducted fan assembly is configured to direct a first portion of airflow to the unducted vane assembly and a second portion of airflow into an inlet duct and to the ducted fan assembly, and wherein the ducted fan assembly is configured to direct the second portion of airflow to a fan duct and to a core duct.

63. A turbomachine engine comprises an open rotor fan assembly, a core engine, a vane assembly, a gearbox, a gearbox efficiency rating. The open rotor fan assembly including a plurality of fan blades. The core engine including a low-pressure compressor, a high-pressure compressor, a combustor, a high-pressure turbine, and a low-pressure turbine. The vane assembly including a plurality of vanes extending radially outwardly from the core engine in a cantilever manner. The gearbox is coupled to the low-pressure turbine and the open rotor fan assembly. The gearbox comprises a gear ratio of 6.0-12.0 and is configured such that a first rotational speed of the open rotor fan assembly is less than a second rotational speed of the low-pressure turbine. The gearbox efficiency rating is 0.18-1.41.

64. The turbomachine engine of any clause herein, wherein a gearbox oil flow rate at an inlet of the gearbox is within a range of 6-36 gallons per minute at a max takeoff condition, wherein a diameter of the fan blades is 140-192 inches, and wherein a net thrust of the turbomachine engine is within a range of 12,000-30,000 pounds force at a max takeoff condition.

65. A turbomachine engine comprises a fan case, a fan assembly, a pitch change mechanism, a core engine, a vane assembly, a gearbox, and a gearbox efficiency rating. The fan assembly is disposed radially within the fan case and comprises a plurality of fan blades. The pitch change mechanism is coupled to the fan assembly and is configured to adjust a pitch of the fan blades. The core engine including a low-pressure turbine. The vane assembly includes a plurality of vanes. The vanes are disposed aft of the fan blades and are coupled to the core engine and the fan case. The gearbox is coupled to the low-pressure turbine and the fan assembly. The gearbox is configured such that a ratio of a first rotational speed of the low-pressure turbine to a second rotational speed of the fan assembly is within a range of 4.1-14.0. The gearbox efficiency rating is 0.10-1.8. The gearbox efficiency rating equals

Q⁡(D1.56T)1.53,
where Q is a gearbox oil flow rate at an inlet of the gearbox measured in gallons per minute at a max takeoff condition, D is a diameter of the fan blades measured in inches, and T is a net thrust of the turbomachine engine measured in pounds force at the max takeoff condition.

66. The turbomachine engine of any clause herein, wherein the pitch change mechanism is a first pitch change mechanism, and wherein the turbomachine engine further comprises a second pitch change mechanism coupled to the vane assembly and configured to adjust a pitch of the vanes.

67. The turbomachine engine of any clause herein, wherein the gearbox efficiency rating is 0.25-1.5.

68. The turbomachine engine of any clause herein, wherein the ratio of the first rotational speed of the low-pressure turbine to the second rotational speed of the fan assembly is within a range of 4.5-12.0.

69. The turbomachine engine of any clause herein, wherein the ratio of the first rotational speed of the low-pressure turbine to the second rotational speed of the fan assembly is within a range of 6.0-11.0.

70. The turbomachine engine of any clause herein, wherein Q is within a range of 5-55 gallons per minute.

71. The turbomachine engine of any clause herein, wherein D is 120-216 inches.

72. The turbomachine engine of any clause herein, wherein T is within a range of 10,000-100,000 pounds force.

73. The turbomachine engine of any clause herein, wherein the gearbox is an epicyclic gearbox comprising a sun gear, a plurality of planet gears, and a ring gear, wherein the sun gear is coupled to the low-pressure turbine, and wherein the ring gear is coupled to the fan assembly.

74. The turbomachine engine of any clause herein, wherein the gearbox is an epicyclic gearbox comprising a sun gear, a plurality of planet gears, and a ring gear, wherein the sun gear is coupled to the low-pressure turbine, wherein the planet gears are coupled to a planet carrier, and wherein the planet carrier is coupled to the fan assembly.

75. The turbomachine engine of any clause herein, wherein the gearbox is a multi-stage gearbox.

76. The turbomachine engine of any clause herein, wherein the gearbox is a compound gearbox.

77. A turbomachine engine comprises a fan case, a fan assembly, a pitch change mechanism, a vane assembly, a core engine, a gearbox, and a gearbox efficiency rating. The fan assembly is housed within the fan case and comprising a plurality of fan blades. The pitch change mechanism is coupled to the fan assembly. The vane assembly is housed within the fan case and comprises a plurality of vanes. The core engine includes a low-pressure compressor, a high-pressure compressor, a combustor, a high-pressure turbine, and a low-pressure turbine. The gearbox is coupled to the low-pressure turbine and the fan assembly. The gearbox comprises a gear ratio of 4.1-14.0 and is configured such that the fan assembly rotates slower than one or more stages of the low-pressure turbine. The gearbox efficiency rating is 0.10-1.8. The gearbox efficiency rating equals

Q⁡(D1.56T)1.53,
where Q is a gearbox oil flow rate at an inlet of the gearbox measured in gallons per minute at a max takeoff condition, D is a diameter of the fan blades measured in inches, and T is a net thrust of the turbomachine engine measured in pounds force at the max takeoff condition.

77. The turbomachine engine of any clause herein, wherein the gearbox efficiency rating is 0.10-1.01.

78. The turbomachine engine of any clause herein, wherein the gearbox efficiency rating is 0.19-1.8

79. The turbomachine engine of any clause herein, wherein Q is within a range of 6-36 gallons per minute.

80. The turbomachine engine of any clause herein, wherein D is 140-192 inches.

81. The turbomachine engine of any clause herein, wherein T is within a range of 10,000-40,000 pounds force.

82. The turbomachine engine of any clause herein, wherein the gearbox is an epicyclic gearbox comprising a star gear configuration.

83. The turbomachine engine of any clause herein, wherein the gearbox is an epicyclic gearbox comprising a planet gear configuration.

85. The turbomachine engine of any clause herein, wherein the low-pressure compressor comprises 1-2 stages.

86. The turbomachine engine of any clause herein, wherein the high-pressure compressor comprises 10-11 stages.

87. The turbomachine engine of any clause herein, wherein the high-pressure turbine comprises two stages.

88. The turbomachine engine of any clause herein, wherein the low-pressure turbine comprises 6-7 stages.

89. The turbomachine engine of any clause herein, wherein the low-pressure turbine is a counter-rotating low-pressure turbine comprising inner blade stages and outer blade stages, wherein the inner blade stages extend radially outwardly from an inner shaft, wherein the outer blade stages extend radially inwardly from an outer drum, and wherein the gearbox is configured such that the fan assembly rotates slower than the inner blade stages of the low-pressure turbine.

90. A turbomachine engine comprises a fan case, a fan assembly, a pitch change mechanism, a core engine, a vane assembly, a gearbox, and a gearbox efficiency rating. The fan assembly includes a plurality of fan blades. The pitch change mechanism is coupled to the fan assembly. The core engine includes a low-pressure compressor, a high-pressure compressor, a combustor, a high-pressure turbine, and a low-pressure turbine. The vane assembly includes a plurality of vanes. The gearbox is coupled to the low-pressure turbine and the fan assembly. The gearbox comprises a gear ratio of 6.0-12.0 and is configured such that a first rotational speed of the fan assembly is less than a second rotational speed of the low-pressure turbine. The gearbox efficiency rating of 0.18-1.41.

91. The turbomachine engine of any clause herein, wherein a gearbox oil flow rate at an inlet of the gearbox is within a range of 5-40 gallons per minute at a max takeoff condition, wherein a diameter of the fan blades is 140-192 inches, and wherein a net thrust of the turbomachine engine is within a range of 12,000-30,000 pounds force at a max takeoff condition.

92. A turbomachine engine comprises a fan case, a fan assembly, a vane assembly, a core engine, a gearbox, and a gearbox efficiency rating. The fan assembly comprises a plurality of fan blades. The vane assembly includes a plurality of vanes, and the vanes are disposed aft of the fan blades. The core engine includes a counter-rotating low-pressure turbine. The gearbox is coupled to the counter-rotating low-pressure turbine and the fan assembly. The gearbox is configured such that a ratio of a first rotational speed of the counter-rotating low-pressure turbine to a second rotational speed of the fan assembly is within a range of 4.1-14.0. The gearbox efficiency rating is 0.10-1.8. The gearbox efficiency rating equals

Q⁡(D1.56T)1.53,
where Q is a gearbox oil flow rate at an inlet of the gearbox measured in gallons per minute at a max takeoff condition, D is a diameter of the fan blades measured in inches, and T is a net thrust of the turbomachine engine measured in pounds force at the max takeoff condition.

93. The turbomachine engine of any clause herein, wherein the gearbox efficiency rating is 0.25-1.15.

94. The turbomachine engine of any clause herein, wherein the ratio of the first rotational speed of the counter-rotating low-pressure turbine to the second rotational speed of the fan assembly is within a range of 4.5-12.0.

95. The turbomachine engine of any clause herein, wherein the ratio of the first rotational speed of the counter-rotating low-pressure turbine to the second rotational speed of the fan assembly is within a range of 6.0-11.0.

96. The turbomachine engine of any clause herein, wherein Q is within a range of 5-55 gallons per minute.

97. The turbomachine engine of any clause herein, wherein D is 120-216 inches.

98. The turbomachine engine of any clause herein, wherein T is within a range of 10,000-100,000 pounds force.

99. The turbomachine engine of any clause herein, wherein the counter-rotating low-pressure turbine includes an inner rotor and an outer drum, wherein the inner rotor comprises a plurality of inner blade stages, wherein the outer drum comprises a plurality of outer blade stages, and wherein the outer blade stages are disposed between adjacent inner blade stages.

100. The turbomachine engine of any clause herein, wherein the counter-rotating low-pressure turbine comprises exactly three inner blade stages and exactly three outer blade stages.

101. The turbomachine engine of any clause herein, wherein the counter-rotating low-pressure turbine comprises exactly four inner blade stages and exactly three outer blade stages.

102. The turbomachine engine of any clause herein, wherein the gearbox is an epicyclic gearbox comprising a sun gear, a plurality of planet gears, and a ring gear, wherein the sun gear is coupled to the inner rotor of the counter-rotating low-pressure turbine, and wherein the ring gear is coupled to outer drum of the counter-rotating low-pressure turbine and the fan assembly.

103. The turbomachine engine of any clause herein, wherein the gearbox is a multi-stage gearbox comprising a first stage and a second stage, wherein the first stage of the gearbox comprises a first-stage sun gear, a plurality of first-stage planet gears coupled to a first-stage planet carrier, and a first-stage ring gear, wherein the second stage of the gearbox comprises a second-stage sun gear, a plurality of second-stage planet gears coupled to a second-stage planet carrier, and a second-stage ring gear, wherein the first-stage sun gear is coupled to the inner rotor of the counter-rotating low-pressure turbine, and wherein second-stage sun gear is coupled to the outer drum of the counter-rotating low-pressure turbine.

104. The turbomachine engine of any clause herein, wherein the first stage of the gearbox comprises a star gear configuration, and wherein the second stage of the gearbox comprises a planet gear configuration.

105. The turbomachine engine of any clause herein, further comprising a pitch change mechanism coupled to the fan assembly and configured to adjust a pitch of the fan blades.

106. A turbomachine engine comprises a fan case, a fan assembly, a vane assembly, a core engine, a gearbox, and a gearbox efficiency rating. The fan assembly is housed within the fan case and comprises a plurality of fan blades. The vane assembly is housed within the fan case and comprises a plurality of vanes. The core engine includes a low-pressure compressor, a high-pressure compressor, a combustor, a high-pressure turbine, and a counter-rotating low-pressure turbine. The gearbox is coupled to the counter-rotating low-pressure turbine and the fan assembly. The gearbox comprises a gear ratio of 4.1-14.0 and is configured such that the fan assembly rotates slower than one or more stages of the counter-rotating low-pressure turbine. The gearbox efficiency rating is 0.10-1.8. The gearbox efficiency rating equals

Q⁡(D1.56T)1.53,
where Q is a gearbox oil flow rate at an inlet of the gearbox measured in gallons per minute at a max takeoff condition, D is a diameter of the fan blades measured in inches, and T is a net thrust of the turbomachine engine measured in pounds force at the max takeoff condition.

107. The turbomachine engine of any clause herein, wherein the gearbox is an epicyclic gearbox comprising a star gear configuration.

108. The turbomachine engine of any clause herein, wherein the gearbox comprises a first stage and a second stage, and wherein the fan assembly rotates slower than all stages of the counter-rotating low-pressure turbine.

109. The turbomachine engine of any clause herein, wherein the first stage of the gearbox comprises a star gear configuration, and wherein the second stage of the gearbox comprises a planet gear configuration.

110. The turbomachine engine of any clause herein, wherein the gearbox efficiency rating is 0.20-1.10.

111. The turbomachine engine of any clause herein, wherein the gearbox efficiency rating is 0.10-1.01.

112. The turbomachine engine of any clause herein, wherein the gearbox efficiency rating is 0.19-1.8.

113. The turbomachine engine of any clause herein, wherein Q is within a range of 6-36 gallons per minute.

114. The turbomachine engine of any clause herein, wherein D is 140-192 inches.

115. The turbomachine engine of any clause herein, wherein T is within a range of 10,000-40,000 pounds force.

117. The turbomachine engine of any clause herein, wherein the low-pressure compressor comprises 1-2 stages.

118. The turbomachine engine of any clause herein, wherein the high-pressure compressor comprises 10-11 stages.

119. The turbomachine engine of any clause herein, wherein the high-pressure turbine comprises two stages.

120. The turbomachine engine of any clause herein, wherein the counter-rotating low-pressure turbine comprises 6-7 stages.

121. The turbomachine engine of any clause herein, wherein the counter-rotating low-pressure turbine comprises inner blade stages and outer blade stages, wherein the inner blades stages are coupled to a first rotatable shaft, and wherein the outer blades stages are coupled to a second rotatable shaft.

122. The turbomachine engine of any clause herein, wherein the gearbox is located forward from the combustor.

123. The turbomachine engine of any clause herein, wherein the gearbox is located aft of the combustor.

124. A turbomachine engine comprises a ducted fan assembly, a pitch change mechanism, a core engine, a ducted vane assembly, a gearbox, and a gearbox efficiency rating. The ducted fan assembly includes a plurality of fan blades. The pitch change mechanism is coupled to the ducted fan assembly. The core engine includes a low-pressure compressor, a high-pressure compressor, a combustor, a high-pressure turbine, and a counter-rotating low-pressure turbine. The ducted vane assembly includes a plurality of vanes. The gearbox is coupled to the counter-rotating low-pressure turbine and the ducted fan assembly. The gearbox comprises a gear ratio of 6.0-12.0 and is configured such that a first rotational speed of the ducted fan assembly is less than a second rotational speed of one or more stages of the counter-rotating low-pressure turbine. The gearbox efficiency rating is 0.18-1.41 at a max takeoff condition.

125. The turbomachine engine of any clause herein, wherein a gearbox oil flow rate at an inlet of the gearbox is within a range of 6-36 gallons per minute at a max takeoff condition, wherein a diameter of the fan blades is 140-192 inches, and wherein a net thrust of the turbomachine engine is within a range of 12,000-30,000 pounds force at a max takeoff condition.

126. A turbomachine engine comprises a fan case, a fan assembly, a vane assembly, a core engine, a gearbox, and a gearbox efficiency rating. The fan assembly is disposed radially within the fan case and comprises a plurality of fan blades. The core engine includes a low-pressure turbine. The vane assembly includes a plurality of vanes, and the vanes are disposed aft of the fan blades and are coupled to the core engine and the fan case. The gearbox is coupled to the low-pressure turbine and the fan assembly, and the gearbox comprises a gear ratio within a range of 4.1-14.0. The gearbox efficiency rating is 0.10-1.8 at a max takeoff condition.

127. The turbomachine engine of any clause herein, wherein the gearbox efficiency rating is 0.25-1.15 at the max takeoff condition.

128. The turbomachine engine of any clause herein, wherein the gearbox efficiency rating is 0.10-1.01 at the max takeoff condition.

129. The turbomachine engine of any clause herein, wherein the gearbox efficiency rating is 0.19-1.8 at the max takeoff condition.

130. The turbomachine engine of any clause herein, wherein the gear ratio of the gearbox is within a range of 4.5-12.0.

131. The turbomachine engine of any clause herein, wherein the gear ratio of the gearbox is within a range of 6.0-11.0.

132. The turbomachine engine of any clause herein, wherein a gearbox oil flow rate at an inlet of the gearbox is within a range of 5-55 gallons per minute at the max takeoff condition.

133. The turbomachine engine of any clause herein, wherein a diameter of the fan blades is 72-216 inches.

134. The turbomachine engine of any clause herein, wherein a net thrust of the turbomachine engine is within a range of 10,000-100,000 pounds force at the max takeoff condition.

135. The turbomachine engine of any clause herein, wherein the gearbox is an epicyclic gearbox comprising a sun gear, a plurality of planet gears, and a ring gear, wherein the sun gear is coupled to the low-pressure turbine, and wherein the ring gear is coupled to the fan assembly.

136. The turbomachine engine of any clause herein, wherein the gearbox is an epicyclic gearbox comprising a sun gear, a plurality of planet gears, and a ring gear, wherein the sun gear is coupled to the low-pressure turbine, wherein the planet gears are coupled to a planet carrier, and wherein the planet carrier is coupled to the fan assembly.

137. The turbomachine engine of any clause herein, wherein the gearbox is a multi-stage gearbox.

138. The turbomachine engine of any clause herein, wherein the gearbox comprises one or more compound gears, wherein each compound gear includes a first portion having a first diameter and a second portion having a second diameter, the second diameter being less than the first diameter.

139. The turbomachine engine of any clause herein, further comprising one or more pitch change mechanisms coupled to the fan assembly or the vane assembly.

140. A turbomachine engine comprises a fan case, a fan assembly, a vane assembly, a core engine, a gearbox, and a gearbox efficiency rating. The fan assembly is housed within the fan case and comprising a plurality of fan blades. The vane assembly is housed within the fan case and comprising a plurality of vanes. The core engine includes a low-pressure compressor, a high-pressure compressor, a combustor, a high-pressure turbine, and a low-pressure turbine. The gearbox is coupled to the low-pressure turbine and the fan assembly. The gearbox comprises a gear ratio of 4.1-14.0 and is configured such that the fan assembly rotates slower than one or more stages of the low-pressure turbine. The gearbox efficiency rating of 0.10-1.8 at a max takeoff condition.

141. The turbomachine engine of any clause herein, wherein the gearbox efficiency rating is 0.20-1.15 at the max takeoff condition.

142. The turbomachine engine of any clause herein, wherein a gearbox oil flow rate at an inlet of the gearbox is within a range of 6-36 gallons per minute at the max takeoff condition.

143. The turbomachine engine of any clause herein, wherein the fan blades comprise a diameter within a range of 72-120 inches.

144. The turbomachine engine of any clause herein, wherein a net thrust of the turbomachine engine is within a range of 10,000-40,000 pounds force at the max takeoff condition.

145. The turbomachine engine of any clause herein, wherein the gearbox is an epicyclic gearbox comprising a star gear configuration.

146. The turbomachine engine of any clause herein, wherein the gearbox is an epicyclic gearbox comprising a planet gear configuration.

148. The turbomachine engine of any clause herein, wherein the low-pressure compressor comprises 3-8 stages.

149. The turbomachine engine of any clause herein, wherein the high-pressure compressor comprises 8-15 stages.

150. The turbomachine engine of any clause herein, wherein the high-pressure turbine comprises 1-2 stages.

151. The turbomachine engine of any clause herein, wherein the low-pressure turbine comprises 3-6 stages.

152. The turbomachine engine of any clause herein, wherein the low-pressure turbine is a counter-rotating low-pressure turbine.

153. A turbomachine engine comprises a fan case, a fan assembly, a vane assembly, a core engine, an epicyclic gearbox, and a gearbox efficiency rating. The fan assembly is housed within the fan case and comprises 16-20 fan blades. The vane assembly is housed within the fan case and comprises a plurality of vanes. The core engine includes a low-pressure compressor, a high-pressure compressor, a combustor, a high-pressure turbine, and a low-pressure turbine. The low-pressure compressor comprises 2-4 stages, the high-pressure compressor comprises 8-10 stages, the high-pressure turbine comprises two stages, and the low-pressure turbine comprises 3-4 stages. The epicyclic gearbox is coupled to the low-pressure turbine and the fan assembly. The epicyclic gearbox comprises a gear ratio of 4.1-14.0 and is configured such that the fan assembly rotates slower than one or more stages of the low-pressure turbine. The gearbox efficiency rating is 0.10-1.8 at a max takeoff condition.

154. The turbomachine engine of any clause herein, wherein the gearbox efficiency rating is 0.25-0.55 at the max takeoff condition.

155. The turbomachine engine of any clause herein, wherein the gear ratio of the epicyclic gearbox is within a range of 6.0-12.0.

156. The turbomachine engine of any clause herein, wherein a gearbox oil flow rate at an inlet of the epicyclic gearbox is within a range of 5-40 gallons per minute at the max takeoff condition.

157. The turbomachine engine of any clause herein, wherein a diameter of the fan blades is 72-120 inches.

158. The turbomachine engine of any clause herein, wherein a net thrust of the turbomachine engine is within a range of 10,000-40,000 pounds force at the max takeoff condition.

159. The turbomachine engine of any clause herein, wherein the epicyclic gearbox comprises a sun gear, a plurality of planet gears, and a ring gear, wherein the sun gear is coupled to the low-pressure turbine, and wherein the ring gear is coupled to the fan assembly.

160. The turbomachine engine of any clause herein, wherein the epicyclic gearbox comprises a sun gear, a plurality of planet gears, and a ring gear, wherein the sun gear is coupled to the low-pressure turbine, wherein the planet gears are coupled to a planet carrier, and wherein the planet carrier is coupled to the fan assembly.

161. The turbomachine engine of any clause herein, wherein the epicyclic gearbox is a multi-stage gearbox.

162. The turbomachine engine of any clause herein, further comprising one or more pitch change mechanisms coupled to the fan assembly or the vane assembly.

163. The turbomachine engine of any clause herein, wherein the gearbox is a high power gearbox.