Patent ID: 12221930

DETAILED DESCRIPTION

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

The word “exemplary” is used herein to mean “serving as an example, instance, or illustration.” Any implementation described herein as “exemplary” is not necessarily to be construed as preferred or advantageous over other implementations.

As used herein, the terms “first,” “second,” and “third” may be used interchangeably to distinguish one component from another and are not intended to signify the location or importance of the individual components.

The terms “forward” and “aft” refer to relative positions within a gas turbine engine or vehicle and refer to the normal operational attitude of the gas turbine engine or vehicle. For example, with regard to a gas turbine engine, forward refers to a position closer to an engine inlet and aft refers to a position closer to an engine nozzle or exhaust.

The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows.

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

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

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

One or more components of the turbomachinery engine or gear assembly described herein below 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 components 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 components 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.

Rising fuel prices, depleting natural resources, and regulatory constraints place increasing demands on turbomachinery engines. As such, turbomachinery engines with improved efficiency and performance are desired. Designing turbomachinery engines, however, is complex, time consuming, and expensive. There are many engine components and parameters to consider (each of various weight), and many are of the components and parameters are interdependent. Therefore, changing one component or one parameter can often create cascading effects requiring one or more other parameters or components to be reconfigured.

Various turbomachinery engines and gear assemblies are disclosed herein. The disclosed turbomachinery engines have improved efficiency and/or performance than typical turbomachinery engines.

The disclosed turbomachinery engines comprise a gearbox and a turbine (e.g., a low-pressure turbine) coupled to the gearbox. The disclosed turbomachinery engines are characterized or defined by one or more parameters of a turbine (e.g., the low-pressure turbine). These turbine parameters include: an area ratio and/or an area-EGT ratio. Additional information about these ratios and exemplary engines comprising these ratios are provided below.

Referring now to the drawings,FIG.1is an example of an engine100including a gear assembly102according to aspects of the present disclosure. The engine100includes a fan assembly104driven by a core engine106. In various examples, 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 examples, the fan assembly104includes eight (8) to twenty-two (22) fan blades108. In particular examples, the fan assembly104includes ten (10) to eighteen (18) fan blades108. In certain examples, the fan assembly104includes twelve (12) to sixteen (16) fan blades108. In some examples, the vane assembly110includes three (3) to thirty (30) vanes112. In certain examples, the vane assembly110includes an equal or fewer quantity of vanes112to fan blades108. For example, in particular examples, the engine100includes twelve (12) fan blades108and ten (10) vanes112. In other examples, the vane assembly110includes a greater quantity of vanes112to fan blades108. For example, in particular implementations, the engine100includes ten (10) fan blades108and twenty-three (23) vanes112.

In certain examples, such as depicted inFIG.1, the vane assembly110is positioned downstream or aft of the fan assembly104. However, it should be appreciated that in some examples, the vane assembly110may be positioned upstream or forward of the fan assembly104. In still various examples, 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 examples, 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 examples, the engine100may be configured as an unshrouded turbofan engine, an open rotor engine, or a propfan engine. In particular examples, 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 Dan is shown (i.e., the radius of the fan). In some examples, the Dfanis 72-216 inches. In particular examples the Dfanis 100-200 inches. In certain examples, the Dfanis 120-190 inches. In other examples, the Dfanis 72-120 inches. In yet other examples, the Dfanis 50-80 inches.

In some examples, the fan blade tip speed at a cruise flight condition can be 650 to 1000 fps, or 800 to 900 fps. A fan pressure ratio (FPR) for the fan assembly104can be 1.04 to 1.10, or in some examples 1.05 to 1.08, as measured across the fan blades at a cruise flight condition. In other examples, the FPR can be within a range of 1.04-1.8, 1.1-1.4, 1.3-1.6, or 1.5-1.8.

Cruise altitude is generally an altitude at which an aircraft levels after climb and prior to descending to an approach flight phase. In various examples, the engine is applied to a vehicle with a cruise altitude up to approximately 65,000 ft. In certain examples, cruise altitude is from approximately 28,000 ft. to approximately 45,000 ft. In still certain examples, cruise altitude is expressed in flight levels (FL) based on standard air pressure at sea level, in which a cruise flight condition is from FL280 to FL650. In another example, cruise flight condition is from FL280 to FL450. In still certain examples, cruise altitude is defined based at least on barometric pressure, in which cruise altitude is from approximately 4.85 psia to 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 example, cruise altitude is from approximately 4.85 psia to approximately 2.14 psia. It should be appreciated that in certain examples, 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 examples, the engine100includes a length (L) from a longitudinally (or axial) forward end116to a longitudinally aft end118. In various examples, the engine100defines a ratio of L/Dcorethat provides for reduced installed drag. In one example, L/Dcoreis at least 2. In another example, L/Dcoreis at least 2.5. In some examples, the L/Dcoreis less than 5, less than 4, and less than 3. In various examples, 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 installed drag may provide for cruise altitude engine and aircraft operation at or above Mach 0.5. In certain examples, 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 from approximately Mach 0.55 to approximately Mach 0.85; or from approximately Mach 0.72 to Mach 0.85 or from approximately Mach 0.75 to Mach 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 assembly102can 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 assembly104and can have a second rotational speed. In some examples, a gear ratio of the first rotational speed to the second rotational speed is less than or equal to four (e.g., within a range of 2.0-4.0). In other examples, a gear ratio of the first rotational speed to the second rotational speed is greater than four (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 configurations. For example, in some instances, 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 examples, the gearbox is a single-stage gearbox (e.g.,FIGS.10-11). In other examples, the gearbox is a multi-stage gearbox (e.g.,FIGS.9and12). In some examples, the gearbox is an epicyclic gearbox. In some examples, 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 examples, a gear ratio of the input rotational speed to the output rotational speed is within a range of 2-4. For example, the gear ratio can be 2-2.9, 3.2-4, or 3.25-3.75). In some examples, a gear ratio of the input rotational speed to the output rotational speed is greater than 4.1. For example, in particular instances, 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 examples, the gear ratio is within a range of 4.5-12 or within a range of 6.0-11.0. As such, in some examples, the fan assembly can be configured to rotate at a rotational speed of 800-1500 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 examples, 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 with 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 example of an open rotor propulsion engine. The engine200is generally similar to the engine100, therefore, like parts will be identified with like numerals increased to the200series, with it being understood that the description of the like parts of the engine100applies to the engine200unless otherwise noted. 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-pressure compressor236and a high-speed turbine238operably rotatably coupled together by a high-speed shaft240. The combustion section232is positioned between the high-pressure compressor236and the high-pressure 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 examples, the combustion section232includes an annular combustor, a can combustor, a cannular combustor, a trapped vortex combustor (TVC), or another appropriate combustion system, or combinations thereof.

The core engine206also includes a booster or low-pressure compressor242positioned 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 compressor 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 instances, a low-pressure compressor (which can also be referred to as “a booster”) 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 (including exactly 3, 4, or 5 stages). For example, in certain examples, an engine can comprise a one stage low-pressure compressor, an 11 stage high-pressure compressor, a two stage high-pressure turbine, 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 turbine, 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 turbine, and a three stage low-pressure turbine. As another example, an engine can comprise a four stage low-pressure compressor, a 10 stage high-pressure compressor, a one stage high-pressure turbine, and a three 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 turbine, and a four stage low-pressure turbine. As another example, an engine can comprise a four stage low-pressure compressor, a 10 stage high-pressure compressor, a one stage high-pressure turbine, and a four stage low-pressure turbine. In other examples, an engine can comprise a 1-3 stage low-pressure compressor, an 8-11 stage high-pressure compressor, a 1-2 stage high-pressure turbine, and a 3-5 stage low-pressure turbine. In some examples, an engine can be configured without a low-pressure compressor.

In some examples, 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 examples, 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 examples, the counter-rotating low-pressure turbine comprises four inner blade stages and three outer blade stages, which can 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-pressure turbine244. Examples 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, examples of the gear assembly202provided herein may be suitable within the radial or diametrical constraints of the core engine206within an engine core cowl272.

Various gearbox configurations are depicted schematically inFIGS.16-19. 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 examples, they are not all equidistant from the rotating assembly.

In some examples, vanes212are mounted to a stationary frame and do not rotate relative to the engine centerline axis220but 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 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-pressure turbine244. In operation, the rotating fan blades208are driven by the low-pressure 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 examples, a gear ratio of the input rotational speed to the output rotational speed is greater than or equal to 4.1. In particular examples, 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 examples, the gear ratio is within a range of 4.5-12 or within a range of 6.0-11.0. As such, in some examples, the fan assembly can be configured to rotate at a rotational speed of 800-1500 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 examples, 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 vanes212incorporate 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 examples, 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 examples, 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 example shown inFIG.2, an annular 360-degree inlet258is located between the fan assembly204and the vane assembly210and 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 example 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 a core duct262or 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 the 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 assembly210at 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 engine106.

Thus, in the example, 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 example 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 that 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 variable bleed valves (VBVs), variable stator vanes (VSVs), 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 examples, 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, Slither900, 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.

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. Exhausting the fan duct flow over the engine core cowl272aids in energizing the boundary layer and enabling the option of a steeper nacelle close out angle between a 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 engine core cowl272reduces 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.1and2). 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 vanes212to be varied more easily. For example,FIGS.3-4, discussed in more detail below, disclose examples in which both the fan assembly and vane assembly are shrouded.

Although depicted as an unshrouded or open rotor engine in the examples 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 turbomachinery configurations, including those for acro-propulsion systems. Certain aspects of the disclosure may be applicable to turbofan, turboprop, or turboshaft engines.

FIG.3is a schematic cross-sectional view of a gas turbine engine in accordance with an example of the present disclosure. More particularly, for the example ofFIG.3, the gas turbine engine is a high-bypass turbofan engine300, referred to herein as “turbofan engine300.” As shown inFIG.3, the turbofan engine300defines an axial direction A (extending parallel to a longitudinal axis302or centerline provided for reference) and a radial direction R (extending perpendicular to the axial direction A). In general, the turbofan engine300includes a fan section304and a core engine306disposed downstream from the fan section304. The turbofan 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 example of turbofan engine300. 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 example 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 assembly336. The 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 some instances, 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.

The gear assembly336(which can also be referred to as “a gearbox”) can, in some examples, comprise a gear ratio of less than or equal to ten. For example, the gearbox336can comprise a gear ratio within a range of 2.0-10.0, 2.0-6.0, 2.0-4.0, 2.0-2.9, 3.0-3.5, 3.2-4.0, 3.25-3.75, etc. In one particular example, the gearbox336can comprise a gear ratio of 3.5.

Referring still to the example ofFIG.3, the disk332is covered by a rotatable front nacelle338aerodynamically contoured to promote 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 example 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 turbofan engine300through an associated inlet350of the nacelle340and/or fan section304. As the volume of air348passes across the fan blades330, a first portion of the air348, as indicated by arrows352, is directed or routed into the bypass airflow passage346and a second portion of the air348, as indicated by arrow354, is 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 blades360(e.g., two stage) that 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 blades364(e.g., four stages) that 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.

It should be noted that a high-pressure turbine (e.g., the HP turbine318) can, in some examples, comprise one or two rotating blade stages and that a low-pressure turbine (e.g., LP turbine320) can, in some instances, comprise three, four, five, six, or seven rotating blade stages.

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 turbofan engine300, 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.

FIG.4is a cross-sectional schematic illustration of an example of an engine400that includes a gear assembly402in combination with a ducted fan assembly404and a core engine. However, unlike the open rotor configuration of the engine200ofFIG.2, the ducted fan assembly404and its fan blades408are contained within an annular fan case480(which can also be referred to as “a nacelle”) and a vane assembly410and vanes412extend radially between a fan cowl452(and/or an 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 engine comprises 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 turbine438(e.g., one stage) and a low-pressure turbine444(e.g., three stage). 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.

It should be noted that a high-pressure turbine (e.g., the high-pressure turbine438) can, in some examples, comprise one or two stages and that a low-pressure turbine (e.g., the low-pressure turbine444) can, in some instances, comprise three, four, five, or six rotating blade stages.

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

In some examples, the engine400can comprise a variable fan nozzle. Operationally, the engine400may include a control system that manages the loading of the fan assembly404, 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.

The fans disclosed herein (e.g., the fan assemblies104,204,304, and404) can comprise various materials. For example, in some instances, a fan can comprise a metal alloy. In some instances, the metal alloy can comprise aluminum, lithium, titanium, and/or other suitable metals for fan blades (e.g., the fan blades108,208,330, and408). In some instances, a fan can comprise composite material. In some examples, a fan can comprise a metal alloy core and a composite cover.

The fans disclosed herein (e.g., the fan assemblies104,204,304, and404) can comprise various dimensions. For example, a fan can comprise a diameter (as measured at the tip of the leading edge) within a range of 72-120 inches (6-10 feet). In some instances, a fan can comprise a diameter within a range of 84-120 inches (7-10 feet) or 84-96 inches (7-8 feet).

The fans disclosed herein comprise a solidity. Solidity is based on average blade chord defined as the blade planform area (surface area on one side of a blade) divided by the blade radial span. The solidity is directly proportional to the number of blades and chord length and inversely proportional to the diameter. For purposes of this disclosure, solidity is equal to the average blade chord (C) times the number of fan blades (N) divided by the product of two (2) times pi (π) times a reference radius (R_ref), which herein is a radius equal to 0.75 times a tip radius of a rotor blade (Rt) (i.e., C×N/(2×π×R_ref)). Using this formula, a fan can comprise a solidity from 0.5 to 1.0, or more particularly from 0.6 to 1.0. In other examples, a fan can comprise a solidity from 1.1 to 1.5, or 1.1-1.3 in certain examples. In still other examples, enhanced performance can be observed when the solidity is greater than or equal to 0.8 and less than or equal to 2, greater than or equal to 0.8 and less than or equal to 1.5, greater than or equal to 1 and less than or equal to 2, or greater than or equal to 1.25 and less than or equal to 1.75.

As mentioned above, rising fuel prices, depleting natural resources, and regulatory constraints place increasing demands on turbomachinery engines. As such, turbomachinery engines with improved efficiency and performance are desired. Designing turbomachinery engines, however, is complex, time consuming, and expensive. There are many engine components and parameters to consider (each of various weight), and many are of the components and parameters are interdependent. Therefore, changing one component or one parameter can often create cascading effects requiring one or more other parameters or components to be reconfigured.

Various turbomachinery engines and gear assemblies are disclosed herein. The disclosed turbomachinery engines have improved efficiency and/or performance than typical turbomachinery engines.

The disclosed turbomachinery engines comprise a gearbox and a turbine (e.g., a low-pressure turbine) coupled to the gearbox. The disclosed turbomachinery engines are characterized or defined by one or more parameters of a turbine (e.g., the low-pressure turbine). These turbine parameters include: an area ratio and/or an area-EGT ratio.

After numerous engine designs, the inventors found unexpectedly that engines comprising the area ratio ranges and/or the area-EGT ratio ranges disclosed herein provide a turbomachine engine with improved performance and efficiency.

The low-pressure turbines disclosed herein comprise 3-5 rotating stages and an area ratio within a range of 2.0-6.5. Each rotating stage comprises an annular exit area defined by a tip radius of a trailing edge of any one blade of the rotating stage and a hub radius of the any one blade of the rotating stage at an axial location aligned with the tip radius. The area ratio equals the annular exit area of an aft-most rotating stage divided by the annular exit area of a forward-most rotating stage.

The low-pressure turbines disclosed herein additionally comprise an area-EGT ratio within

ratio=(area⁢ratio)(1/(stages-1))(EGT/1000).
a range of 1.3-1.6. The area-EGT Each rotating stage comprises an annular exit area defined by a tip radius of a trailing edge of any one blade of the rotating stage and a hub radius of the any one blade of the rotating stage at an axial location aligned with the tip radius. The area ratio is the annular exit area of an aft-most rotating stage divided by the annular exit area of a forward-most rotating stage, wherein the stages is the number of rotating stages. The EGT is an exhaust gas temperature measured in degrees Celsius at an inlet of the turbine at a redline operating condition.

In some examples, a turbomachinery engine includes a fan assembly, a low-pressure turbine, and a gearbox. The fan assembly includes a plurality of fan blades. The low-pressure turbine includes 3-5 rotating stages. Each rotating stage of the low-pressure turbine includes an annular exit area defined by a tip radius of a trailing edge of any one blade of the rotating stage and a hub radius of the any one blade of the rotating stage at an axial location aligned with the tip radius. The low-pressure turbine includes an area ratio equal to the annular exit area of an aft-most rotating stage of the low-pressure turbine divided by the annular exit area of a forward-most rotating stage of the low-pressure turbine, and the area ratio is within a range of 2.0-6.5. The gearbox includes an input and an output. The input of the gearbox is coupled to the low-pressure turbine and includes a first rotational speed, and the output of the gearbox is coupled to the fan assembly and has a second rotational speed.

In some examples, the area ratio of the low-pressure turbine is within a range of 2.0-3.2. In some examples, the area ratio of the low-pressure turbine is within a range of 2.2-4.6.

In some instances, the low-pressure turbine includes exactly three rotating stages and/or the area ratio of the low-pressure turbine is within a range of 2.2-2.91.

In some instances, the low-pressure turbine includes exactly four rotating stages, and/or the area ratio of the low-pressure turbine is within a range of 3.1-5.1.

In some instances, the low-pressure turbine includes exactly five rotating stages, and/or the area ratio of the low-pressure turbine is within a range of 5.0-6.5.

In some examples, a turbomachinery engine includes a fan assembly, a low-pressure turbine, and a gearbox. The fan assembly includes a plurality of fan blades. The low-pressure turbine comprising 3-4 rotating stages. Each rotating stage of the low-pressure turbine comprises an annular exit area defined by a tip radius of a trailing edge of any one blade of the rotating stage and a hub radius of the any one blade of the rotating stage at an axial location aligned with the tip radius. The low-pressure turbine comprises an area ratio equal to the annular exit area of an aft-most rotating stage of the low-pressure turbine divided by the annular exit area of a forward-most rotating stage of the low-pressure turbine, and the area ratio is within a range of 2.1-4.6. The gearbox including an input and an output. The input of the gearbox is coupled to the low-pressure turbine and comprises a first rotational speed, the output of the gearbox 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 3.0-3.5.

In some examples, a turbomachinery engine comprises a fan assembly, a low-pressure turbine, and a gearbox. The fan assembly includes a plurality of fan blades. The low-pressure turbine comprises 4-5 rotating stages. Each rotating stage of the low-pressure turbine comprises an annular exit area defined by a tip radius of a trailing edge of any one blade of the rotating stage and a hub radius of the any one blade of the rotating stage at an axial location aligned with the tip radius. The low-pressure turbine comprises an area ratio equal to the annular exit area of an aft-most rotating stage of the low-pressure turbine divided by the annular exit area of a forward-most rotating stage of the low-pressure turbine, and the area ratio is within a range of 3.6-5.79. The gearbox includes an input and an output. The input of the gearbox is coupled to the low-pressure turbine and comprises a first rotational speed, the output of the gearbox 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 5.0-10.0.

In some examples, a turbomachinery engine comprising a fan assembly, a low-pressure turbine, and a gearbox. The fan assembly including a plurality of fan blades. The low-pressure turbine comprises 3-5 rotating stages and an area-EGT ratio within a range of 1.3-1.6. The area-EGT

ratio=(area⁢ratio)(1/(LPT⁢stages-1))(EGT/1000).
Each rotating stage of the low-pressure turbine comprises an annular exit area defined by a tip radius of a trailing edge of any one blade of the rotating stage and a hub radius of the any one blade of the rotating stage at an axial location aligned with the tip radius. The area ratio is the annular exit area of an aft-most rotating stage of the low-pressure turbine divided by the annular exit area of a forward-most rotating stage of the low-pressure turbine. The LPT stages is the number of rotating stages of the low-pressure turbine. The EGT is an exhaust gas temperature of the low-pressure turbine measured in degrees Celsius at an inlet of the low-pressure turbine at a redline operating condition. The gearbox including an input and an output. The input of the gearbox is coupled to the low-pressure turbine and comprises a first rotational speed, and the output of the gearbox is coupled to the fan assembly and has a second rotational speed.

In some examples, the area-EGT ratio is within a range of 1.38-1.58.

In some examples, the area-EGT ratio is within a range of 1.31-1.53.

In some examples, the area-EGT ratio is within a range of 1.30-1.36.

FIG.5Adepicts a portion of a three-stage low-pressure turbine500, according to one example of the disclosed technology. The low-pressure turbine (LPT)500can be used, for example, with any of the turbomachinery engines disclosed herein (e.g., the engines100,200,300, and400), and particularly Engine 04 depicted in the table ofFIG.9. The LPT500comprises a plurality of rotating blade stages and a plurality of stationary vane stages. In particular, the depicted portion of the LPT500comprises three rotating blade stages502a,502b, and502cand two stationary vane stages504aand504b. The rotating blade stages are referred to herein generically or collectively as “a/the rotating blade stage(s)502” or simply “the blades502,” and the stationary vane stages are referred to herein generically or collectively as “a/the stationary vane stage(s)504” or simply “the vanes504.”

The blades502and the vanes504are disposed within a duct506, which guides the fluid flow through the LPT500.

Each rotating blade stage502of the LPT500comprises an annular exit area defined by a tip radius of a trailing edge of any blade of the rotating stage (or a nominal tip radius of the stage) and a hub radius of the blade of the rotating stage (or a nominal hub radius of the stage) at the axial location aligned with the tip radius. With respect shrouded turbine blades, the tip radius is the radius at the tip of the blade portion, excluding the shroud portion.

For example, the forward-most rotating blade stage (which can also be referred to as “the first stage”)502aof the LPT500comprises an annular exit area508a, as depicted inFIG.5B. The annular exit area508ais defined by the tip radius Rtip1and hub radius Rhub1. Rtip1is the tip radius of the trailing edge of any blade of the first stage502a(or a nominal tip radius of the trailing edges of the blades of the first stage502a), and Rhub1is the hub radius of the blade (or a nominal hub radius of the blades of the first stage) at the axial location aligned with the tip radius Rtip1.

In some examples, the annular exit area of the first stage502acan be within a range of 155-380 in2or within a range of 155-372 in2. In particular examples, the annular exit area can be within a range of 280-380 in2or within a range of 285-372 in2. In the depicted example, the annular exit area508aof the first stage502ais about 327 in2. Additional examples of annular exit areas for the first stage of a three-stage low-pressure turbine are provided in the table depicted inFIG.9.

As another example, the second stage502bof the LPT500comprises an annular exit area. The annular exit area of the second stage502bis defined by the tip radius Rtip2and hub radius Rhub2. Rtip2is the tip radius of the trailing edge of any blade of the second stage502b(or a nominal tip radius of the trailing edges of the blades of the second stage502b), and Rhub2is the hub radius of the blade (or a nominal hub radius of the blades of the second stage502b) at the axial location aligned with the tip radius Rtip2.

In some examples, the annular exit area of the second stage502bcan be within a range of 230-750 in2or within a range of 250-700 in2. In particular examples, the annular exit area of the second stage502bcan be within a range of 450-750 in2or within a range of 462-699 in2. In the depicted example, the annular exit area of the second stage502bis about 526 in2. Additional examples of annular exit areas for the second stage of a three-stage low-pressure turbine are provided in the table depicted inFIG.9.

As another example, the aft-most stage (which can also be referred to as “the third stage”)502cof the LPT500comprises an annular exit area. The annular exit area of the third stage502cis defined by the tip radius Rtip3and hub radius Rhub3. Rtip3is the tip radius of the trailing edge of any blade of the third stage502c(or a nominal tip radius of the trailing edges of the blades of the third stage502c), and Rhub3is the hub radius of the blade (or a nominal hub radius of the blades of the third stage502c) at the axial location aligned with the tip radius Rtip3.

In some examples, the annular exit area of the third stage502ccan be within a range of 350-1050 in2or within a range of 379-1027 in2. In particular examples, the annular exit area of the third stage502ccan be within a range of 600-1050 in2or within a range of 639-1027 in2. In the depicted example, the annular exit area of the third stage502cis about 725 in2. Additional examples of annular exit areas for the third stage of a low-pressure turbine are provided in the table depicted inFIG.9.

The LPT500comprises an area ratio (which can also be referred to as “an exit area ratio”) within a range of 2.0-6.5, within a range of 2.0-3.0, within a range of 2.2-2.91, and specifically about 2.2. The area ratio equals the annular exit area of an aft-most rotating stage of the low-pressure turbine divided by the annular exit area of a forward-most rotating stage of the low-pressure turbine. For example, for the LPT500, the area ratio equals the annular exit area of the third stage502cdivided by the annular exit area508aof the first stage502a.

In addition to having an area ratio within a range of 2.0-6.5, the LPT500can comprise an area-exhaust gas temperature (EGT) ratio, referred to herein as area-EGT ratio, within a range of 1.3-1.6, within a range of 1.38-1.58, and specifically about 1.38. The area-EGT ratio is defined according to Expression (1):

area⁢‐⁢EGT⁢ratio=(the⁢area⁢ratio)(1/(LPT⁢stages-1))(EGT/1000)(1)
where the area ratio is as defined above, LPT stages is the number of rotating blade stages of LPT, and EGT is an exhaust gas temperature of the LPT measured in degrees Celsius at an inlet of the LPT at a redline operating condition.

In some examples, the number of LPT stages is 3, 4, or 5. For example, the LPT500includes exactly three stages. The table ofFIG.9provides additional exemplary engines comprising exactly three LPT stages.

In some examples, EGT is within a range of 1060-1180 degrees Celsius measured at the inlet of the LPT at the redline operating condition. For example, the EGT of the LPT500is about 1083 degrees Celsius at the redline operating condition. As used herein the inlet of the LPT is defined by the turbine vane frame (TVF). The EGT can be measured at any axial location aligned with the TVF, i.e., from the leading edge to the trailing edge of the TVF. Thus, with respect to the LPT500, the inlet of the LPT500for purposes of measuring EGT is any axial location aligned with a TVF510.

Additional information about the LPT500, and its corresponding engine (i.e., Engine 04), is provided in the table depicted inFIG.9.

FIG.6depicts a portion of a low-pressure turbine600, according to one example of the disclosed technology. The low-pressure turbine600can be used, for example, with any of the turbomachinery engines disclosed herein (e.g., the engines100,200,300, and400), and particularly Engine 05 depicted in the table ofFIG.9. The LPT600comprises a plurality of rotating blade stages and a plurality of stationary vane stages. In particular, the depicted portion of the LPT600comprises three rotating blade stages602a,602b, and602cand two stationary vane stages604aand604b. The rotating blade stages are referred to herein generically or collectively as “a/the rotating blade stage(s)602” or simply “the blades602,” and the stationary vane stages are referred to herein generically or collectively as “a/the stationary vane stage(s)604” or simply “the vanes604.”

The blades602and the vanes604are disposed within a duct606aft of a TVF610, which guides the fluid flow through the LPT600.

Each rotating blade stage602of the LPT600comprises an annular exit area defined by a tip radius of a trailing edge of any blade of the rotating stage (or a nominal tip radius of the stage) and a hub radius of the blade of the rotating stage (or a nominal hub radius of the stage) at the axial location aligned with the tip radius.

For example, the forward-most rotating blade stage (which can also be referred to as “the first stage”)602aof the LPT600comprises an annular exit area. The annular exit area is defined by the tip radius Rtip1and hub radius Rhub1. Rtip1is the tip radius of the trailing edge of any blade of the first stage602a(or a nominal tip radius of the trailing edges of the blades of the first stage602a), and Rhub1is the hub radius of the blade (or a nominal hub radius of the blades of the first stage) at the axial location aligned with the tip radius Rtip1.

In some examples, the annular exit area of the first stage602acan be within a range of 155-380 in2. In particular examples, the annular exit area can be within a range of 280-380 in2or within a range of 285-372 in2. In the depicted example, the annular exit area of the first stage602ais about 327 in2.

As another example, the second stage602bof the LPT600comprises an annular exit area. The annular exit area of the second stage602bis defined by the tip radius Rtip2and hub radius Rhub2. Rtip2is the tip radius of the trailing edge of any blade of the second stage602b(or a nominal tip radius of the trailing edges of the blades of the second stage602b), and Rhub2is the hub radius of the blade (or a nominal hub radius of the blades of the second stage602b) at the axial location aligned with the tip radius Rtip2.

In some examples, the annular exit area of the second stage602bcan be within a range of 230-750 in2or within a range of 250-700 in2. In particular examples, the annular exit area of the second stage602bcan be within a range of 450-750 in2or within a range of 462-699 in2. In the depicted example, the annular exit area of the second stage602bis about 577 in2.

As another example, the aft-most stage (which can also be referred to as “the third stage”)602cof the LPT600comprises an annular exit area. The annular exit area of the third stage602cis defined by the tip radius Rtip3and hub radius Rhub3. Rtip3is the tip radius of the trailing edge of any blade of the third stage602c(or a nominal tip radius of the trailing edges of the blades of the third stage602c), and Rhub3is the hub radius of the blade (or a nominal hub radius of the blades of the third stage602c) at the axial location aligned with the tip radius Rtip3.

In some examples, the annular exit area of the third stage602ccan be within a range of 350-1050 in2or within a range of 379-1027 in2. In particular examples, the annular exit area of the third stage602ccan be within a range of 700-1050 in2or within a range of 639-1027 in2. In the depicted example, the annular exit area of the third stage602cis about 827 in2.

The LPT600comprises an area ratio (which can also be referred to as “an exit area ratio”) within a range of 2.0-6.5, within a range of 2.0-3.0, within a range of 2.2-2.9, and specifically about 2.53. For example, for the LPT600, the area ratio equals the annular exit area of the third stage602cdivided by the annular exit area of the first stage602a.

The LPT600can also comprise an area-EGT ratio within a range of 1.3-1.6, within a range of 1.35-1.58, and specifically about 1.47.

In some examples, the EGT of the LPT600is within a range of 1060-1180 degrees Celsius measured at the inlet of the LPT at the redline operating condition. For example, the LPT600comprises an EGT of about 1083 degrees Celsius at the redline operating condition.

Additional information about the LPT600, and its corresponding engine (i.e., Engine 05), is provided in the table depicted inFIG.9.

FIG.7depicts a portion of a low-pressure turbine700, according to one example of the disclosed technology. The low-pressure turbine700can be used, for example, with any of the turbomachinery engines disclosed herein (e.g., the engines100,200,300, and400), and particularly Engine 07 depicted in the table ofFIG.9. The LPT700comprises a plurality of rotating blade stages and a plurality of stationary vane stages. In particular, the depicted portion of the LPT700comprises three rotating blade stages702a,702b, and702cand two stationary vane stages704aand704b. The rotating blade stages are referred to herein generically or collectively as “a/the rotating blade stage(s)702” or simply “the blades702,” and the stationary vane stages are referred to herein generically or collectively as “a/the stationary vane stage(s)704” or simply “the vanes704.”

The blades702and the vanes704are disposed within a duct706aft of a TVF710, which guides the fluid flow through the LPT700.

Each rotating blade stage702of the LPT700comprises an annular exit area defined by a tip radius of a trailing edge of any blade of the rotating stage (or a nominal tip radius of the stage) and a hub radius of the blade of the rotating stage (or a nominal hub radius of the stage) at the axial location aligned with the tip radius. With respect shrouded turbine blades, the tip radius is the radius at the tip of the blade portion, excluding the shroud portion.

For example, the forward-most rotating blade stage (which can also be referred to as “the first stage”)702aof the LPT700comprises an annular exit area. The annular exit area is defined by the tip radius Rtip1and hub radius Rhub1. Rtip1is the tip radius of the trailing edge of any blade of the first stage702a(or a nominal tip radius of the trailing edges of the blades of the first stage702a), and Rhub1is the hub radius of the blade (or a nominal hub radius of the blades of the first stage) at the axial location aligned with the tip radius Rtip1.

In some examples, the annular exit area of the first stage702acan be within a range of 155-380 in2. In particular examples, the annular exit area can be within a range of 280-380 in2. In the depicted example, the annular exit area of the first stage702ais about 372 in2.

As another example, the second stage702bof the LPT700comprises an annular exit area. The annular exit area of the second stage702bis defined by the tip radius Rtip2and hub radius Rhub2. Rtip2is the tip radius of the trailing edge of any blade of the second stage702b(or a nominal tip radius of the trailing edges of the blades of the second stage702b), and Rhub2is the hub radius of the blade (or a nominal hub radius of the blades of the second stage702b) at the axial location aligned with the tip radius Rtip2.

In some examples, the annular exit area of the second stage702bcan be within a range of 250-710 in2. In particular examples, the annular exit area of the second stage702bcan be within a range of 450-700 in2. In the depicted example, the annular exit area of the second stage702bis about 700 in2.

As another example, the aft-most stage (which can also be referred to as “the third stage”)702cof the LPT700comprises an annular exit area. The annular exit area of the third stage702cis defined by the tip radius Rtip3and hub radius Rhub3. Rtip3is the tip radius of the trailing edge of any blade of the third stage702c(or a nominal tip radius of the trailing edges of the blades of the third stage702c), and Rhub3is the hub radius of the blade (or a nominal hub radius of the blades of the third stage702c) at the axial location aligned with the tip radius Rtip3.

In some examples, the annular exit area of the third stage702ccan be within a range of 350-1100 in2. In particular examples, the annular exit area of the third stage702ccan be within a range of 380-1050 in2or within a range of 639-1027 in2. In the depicted example, the annular exit area of the third stage702cis about 1027 in2.

The LPT700comprises an area ratio (which can also be referred to as “an exit area ratio”) within a range of 2.0-6.5, within a range of 2.22-2.91, and specifically about 2.76. For example, for the LPT700, the area ratio equals the annular exit area of the third stage702cdivided by the annular exit area of the first stage702a.

The LPT700can, additionally or alternatively to the area ratio within a range of 2.0-6.5, comprise an area-EGT ratio within a range of 1.3-1.6, within a range of 1.35-1.55, and specifically 1.53.

In some examples, EGT of the LPT700is within a range of 1060-1180 degrees Celsius measured at the inlet of the LPT at the redline operating condition. For example, the LPT700comprises an EGT of about 1083 degrees Celsius at the redline operating condition.

Additional information about the LPT700, and its corresponding engine (i.e., Engine 07), is provided in the table depicted inFIG.9.

FIG.8depicts a portion of a low-pressure turbine800, according to one example of the disclosed technology. The low-pressure turbine800can be used, for example, with any of the turbomachinery engines disclosed herein (e.g., the engines100,200,300, and400), and particularly Engine 09 depicted in the table ofFIG.9. The LPT800comprises a plurality of rotating blade stages and a plurality of stationary vane stages. In particular, the depicted portion of the LPT800comprises three rotating blade stages802a,802b, and802cand two stationary vane stages804aand804b. The rotating blade stages are referred to herein generically or collectively as “a/the rotating blade stage(s)802” or simply “the blades802,” and the stationary vane stages are referred to herein generically or collectively as “a/the stationary vane stage(s)804” or simply “the vanes804.”

The blades802and the vanes804are disposed within a duct806aft of a TVF810, which guides the fluid flow through the LPT800.

Each rotating blade stage802of the LPT800comprises an annular exit area defined by a tip radius of a trailing edge of any blade of the rotating stage (or a nominal tip radius of the stage) and a hub radius of the blade of the rotating stage (or a nominal hub radius of the stage) at the axial location aligned with the tip radius. With respect shrouded turbine blades, the tip radius is the radius at the tip of the blade portion, excluding the shroud portion.

The LPT800comprises three rotating blade stages, a redline EGT of 1067 degrees Celsius, a first stage exit area of 293.2 in2, a second stage exit area of476in2, a third stage exit area of 764.9 in2, an area ratio of 2.61, an area-EGT ratio of 1.51, a first stage AN2value of 30, and a third stage AN2value of 80. AN2the product of A and N2, where A is the annular exit area of a particular rotating stage of the low-pressure turbine measured in square inches, N is the rotational speed of the low-pressure turbine measured in revolutions per minute at a redline operating condition, and the product of AN2is divided by 109.

FIG.9provides additional information about the LPT800(see Engine 09).

FIG.9provides a table with several additional examples of turbomachinery engines comprising three rotating blade stages, a LPT with an area ratio within a range of 2.0-6.5 and an area-EGT ratio within a range of 1.3-1.6. The engines disclosed inFIG.9comprise a gear ratio of 2-9 or 2.95-8.33. The EGT at a redline operating condition for the engines ofFIG.9is within a range of 1060-1175 degrees Celsius or within a range of 1067-1083 degrees Celsius. The exit area of stage 1 of the disclosed engines is within a range of 155-380 in2or within a range of 285.0-372.4 in2. The exit area of stage 2 of the engines ofFIG.9is within a range of 230-750 in2or within a range of 461.8-699.5 in2. The exit area of stage 3 of the engines ofFIG.9is within a range of 350-1050 in2or within a range of 638.5-1026.5 in2. The area ratio of the engines disclosed inFIG.9is within a range of 2.0-6.5 or within a range of 2.22-2.91. The area-EGT ratio is within a range of 1.3-1.6 or within a range of 1.38-1.58. The engines disclosed inFIG.9comprise a first stage AN2value within a range of 9-36 or within a range of 30-36 at a redline operating condition. The engines disclosed inFIG.9comprise a third stage (exit) AN2value within a range of 44-104 or within a range of 79-104 at a redline operating condition.

FIGS.10-12provide examples of low-pressure turbines comprising four rotating blade stages. The disclosed LPTs comprise an area ratio within a range of 2.0-6.5 and an area-EGT ratio within a range of 1.3-1.6.

FIG.10depicts a portion of a four-stage low-pressure turbine900, according to one example of the disclosed technology. The low-pressure turbine900can be used, for example, with any of the turbomachinery engines disclosed herein (e.g., the engines100,200,300, and400), and particularly Engine 16 depicted in the table ofFIG.12. The LPT900comprises a plurality of rotating blade stages and a plurality of stationary vane stages. In particular, the depicted portion of the LPT900comprises four rotating blade stages902a,902b,902c, and902dand three stationary vane stages904a,904b, and904c. The rotating blade stages are referred to herein generically or collectively as “a/the rotating blade stage(s)902” or simply “the blades902,” and the stationary vane stages are referred to herein generically or collectively as “a/the stationary vane stage(s)904” or simply “the vanes904.”

The blades902and the vanes904are disposed within a duct906aft of a TVF910, which guides the fluid flow through the LPT900.

Each rotating blade stage902of the LPT900comprises an annular exit area defined by a tip radius of a trailing edge of any blade of the rotating stage (or a nominal tip radius of the stage) and a hub radius of the blade of the rotating stage (or a nominal hub radius of the stage) at the axial location aligned with the tip radius. With respect shrouded turbine blades, the tip radius is the radius at the tip of the blade portion, excluding the shroud portion.

The LPT900comprises four rotating blade stages, a redline EGT of 1080 degrees Celsius, a first stage exit area of 299.2 in2, a second stage exit area of 442.1 in2, a third stage exit area of 618.3 in2, a fourth stage exit area of 998.1 in2, an area ratio of 3.34, an area-EGT ratio of 1.38, a first stage AN2value of 13, and a fourth stage AN2value of 44.FIG.12provides additional information about the LPT900(see Engine 16).

FIG.11depicts a portion of a four-stage low-pressure turbine1000, according to one example of the disclosed technology. The low-pressure turbine1000can be used, for example, with any of the turbomachinery engines disclosed herein (e.g., the engines100,200,300, and400), and particularly Engine 16 depicted in the table ofFIG.12. The LPT1000comprises a plurality of rotating blade stages and a plurality of stationary vane stages. In particular, the depicted portion of the LPT1000comprises four rotating blade stages1002a,1002b,1002c, and1002dand three stationary vane stages1004a,1004b, and1004c. The rotating blade stages are referred to herein generically or collectively as “a/the rotating blade stage(s)1002” or simply “the blades1002,” and the stationary vane stages are referred to herein generically or collectively as “a/the stationary vane stage(s)1004” or simply “the vanes1004.”

The blades1002and the vanes1004are disposed within a duct1006aft of a TVF1010, which guides the fluid flow through the LPT1000.

Each rotating blade stage1002of the LPT1000comprises an annular exit area defined by a tip radius of a trailing edge of any blade of the rotating stage (or a nominal tip radius of the stage) and a hub radius of the blade of the rotating stage (or a nominal hub radius of the stage) at the axial location aligned with the tip radius. With respect shrouded turbine blades, the tip radius is the radius at the tip of the blade portion, excluding the shroud portion.

The LPT1000comprises four rotating blade stages, a redline EGT of 1175 degrees Celsius, a first stage exit area of 222.4 in2, a second stage exit area of 350.7 in2, a third stage exit area of 612.1 in2, a fourth stage exit area of 907.7 in2, an area ratio of 4.08, an area-EGT ratio of 1.36, a first stage AN2value of 20, and a fourth stage AN2value of 80.FIG.12provides additional information about the LPT1000(see Engine 22).

FIG.12provides a table with several additional examples of turbomachinery engines comprising four rotating blade stages, a LPT with an area ratio within a range of 2.0-6.5 and an area-EGT ratio within a range of 1.3-1.6. The engines disclosed inFIG.12comprise a gear ratio of 2.0-9.0 or 2.33-8.70. The EGT at a redline operating condition for the engines ofFIG.12is within a range of 1060-1175 degrees Celsius. The exit area of stage 1 of the disclosed engines is within a range of 155-380 in2or within a range of 171.9-299.2 in2. The exit area of stage 2 of the engines ofFIG.12is within a range of 230-750 in2or within a range of 275.0-516.4 in2. The exit area of stage 3 of the engines ofFIG.12is within a range of 350-1050 in2or within a range of 496.9-850.8 in2. The exit area of stage 4 of the engines ofFIG.12is within a range of 630-1200 in2, within a range of 632-1186 in2, or within a range of 632.0-1185.1 in2. The area ratio of the engines disclosed inFIG.12is within a range of 2.0-6.5 or within a range of 3.06-5.09. The area-EGT ratio is within a range of 1.3-1.6 or within a range of 1.36-1.53. The engines disclosed inFIG.12comprise a first stage AN2value within a range of 9-36 or within a range of 13-24 at a redline operating condition. The engines disclosed inFIG.12comprise a fourth stage (exit) AN2value within a range of 44-104 or within a range of 44-94 at a redline operating condition.

FIGS.13-15provide examples of low-pressure turbines comprises five rotating blade stages. The disclosed LPTs comprise an area ratio within a range of 2.0-6.5 and an area-EGT ratio within a range of 1.3-1.6.

FIG.13depicts a portion of a five-stage low-pressure turbine1100, according to one example of the disclosed technology. The low-pressure turbine1100can be used, for example, with any of the turbomachinery engines disclosed herein (e.g., the engines100,200,300, and400), and particularly Engine 33 depicted in the table ofFIG.15. The LPT1100comprises a plurality of rotating blade stages and a plurality of stationary vane stages. In particular, the depicted portion of the LPT1100comprises five rotating blade stages1102a,1102b,1102c,1102d, and1102eand four stationary vane stages1104a,1104b,1104c, and1104d. The rotating blade stages are referred to herein generically or collectively as “a/the rotating blade stage(s)1102” or simply “the blades1102,” and the stationary vane stages are referred to herein generically or collectively as “a/the stationary vane stage(s)1104” or simply “the vanes1104.”

The blades1102and the vanes1104are disposed within a duct1106aft of a TVF1110, which guides the fluid flow through the LPT1100.

Each rotating blade stage1102of the LPT1100comprises an annular exit area defined by a tip radius of a trailing edge of any blade of the rotating stage (or a nominal tip radius of the stage) and a hub radius of the blade of the rotating stage (or a nominal hub radius of the stage) at the axial location aligned with the tip radius. With respect shrouded turbine blades, the tip radius is the radius at the tip of the blade portion, excluding the shroud portion.

The LPT1100comprises five rotating blade stages, a redline EGT of 1175 degrees Celsius, a first stage exit area of 212.1 in2, a second stage exit area of 341.6 in2, a third stage exit area of 524.5 in2, a fourth stage exit area of 875.0 in2, a fifth stage exit area of 1212.0 in2, an area ratio of 5.72, an area-EGT ratio of 1.32, a first stage AN2value of 15, and a fifth stage AN2value of 84.FIG.15provides additional information about the LPT1100(see Engine 33).

FIG.14depicts a portion of a five-stage low-pressure turbine1200, according to one example of the disclosed technology. The low-pressure turbine1200can be used, for example, with any of the turbomachinery engines disclosed herein (e.g., the engines100,200,300, and400), and particularly Engine 36 depicted in the table ofFIG.15. The LPT1200comprises a plurality of rotating blade stages and a plurality of stationary vane stages. In particular, the depicted portion of the LPT1200comprises five rotating blade stages1202a,1202b,1202c,1202d, and1202eand four stationary vane stages1204a,1204b,1204c, and1204d. The rotating blade stages are referred to herein generically or collectively as “a/the rotating blade stage(s)1202” or simply “the blades1202,” and the stationary vane stages are referred to herein generically or collectively as “a/the stationary vane stage(s)1204” or simply “the vanes1204.”

The blades1202and the vanes1204are disposed within a duct1206aft of a TVF1210, which guides the fluid flow through the LPT1200.

Each rotating blade stage1202of the LPT1200comprises an annular exit area defined by a tip radius of a trailing edge of any blade of the rotating stage (or a nominal tip radius of the stage) and a hub radius of the blade of the rotating stage (or a nominal hub radius of the stage) at the axial location aligned with the tip radius. With respect shrouded turbine blades, the tip radius is the radius at the tip of the blade portion, excluding the shroud portion.

The LPT1200comprises five rotating blade stages, a redline EGT of 1175 degrees Celsius, a first stage exit area of 232.6 in2, a second stage exit area of 326.9 in2, a third stage exit area of 527.7 in2, a fourth stage exit area of 895.0 in2, a fifth stage exit area of 1279.3 in2, an area ratio of 5.5, an area-EGT ratio of 1.30, a first stage AN2value of 14, and a fifth stage AN2value of 76.FIG.15provides additional information about the LPT1200(see Engine 36).

FIG.15provides a table with several additional examples of turbomachinery engines comprising five rotating blade stages, a LPT with an area ratio within a range of 2.0-6.5 and an area-EGT ratio within a range of 1.3-1.6. The engines disclosed inFIG.15comprise a gear ratio of 2.0-9.0 or 6.96-7.56. The EGT at a redline operating condition for the engines ofFIG.15is within a range of 1060-1175 degrees Celsius, and particularly 1175 degrees Celsius. The exit area of stage 1 of the disclosed engines is within a range of 155-380 in2or within a range of 155.4-232.6 in2. The exit area of stage 2 of the engines ofFIG.15is within a range of 230-750 in2or within a range of 250.2-341.6 in2. The exit area of stage 3 of the engines ofFIG.15is within a range of 350-1050 in2or within a range of 379.3-527.7 in2. The exit area of stage 4 of the engines ofFIG.15is within a range of 630-1200 in2or within a range of 563.1-895 in2. The exit area of stage 5 of the engines ofFIG.15is within a range of 800-1300 in2, within a range of 851-1280 in2, or within a range of 851.4-1279.3 in2. The area ratio of the engines disclosed inFIG.15is within a range of 2.0-6.5 or within a range of 5.48-6.43. The area-EGT ratio is within a range of 1.3-1.6 or within a range of 1.30-1.36. The engines disclosed inFIG.15comprise a first stage AN2value within a range of 9-36 or within a range of 9-15 at a redline operating condition. The engines disclosed inFIG.15comprise a fifth stage (exit) AN2value within a range of 44-104 or within a range of 59-84 at a redline operating condition.

The low-pressure turbines disclosed herein comprising an area ratio within a range of 2.0-6.5 and an area-EGT ratio within a range of 1.3-1.6 provides one or more advantages over conventional low-pressure turbines. In some examples, the disclosed LPTs have up to +1.3% (e.g., +0.1% to +1.3%) LPT efficiency compared to conventional LPTs. In some examples, the disclosed LPTs enable reduced LPT stage count or reduced tip speeds, which provides weight and/or cost reduction, without an efficiency penalty. In some examples, the disclosed LPTs enable higher BPR engines without adding LPT stages. In some examples, the disclosed LPTs reduce turbine rear frame (TRF) loss by up to 0.3% dP/P1due to the reduced LPT exit Mach number. As used herein, “dP” is the change in fluid pressure across the TRF, and “P1” is the fluid pressure prior to the TRF. Stated another way, dP/P1equals the fluid pressure after the TRF (P2) minus the fluid pressure prior to the TRF (P1) divided by P1. Thus, dP/P1is the relative change of the fluid pressure across the TRF. In at least some instances, the LPT exit Mach number of the LPTs disclosed herein can be <0.48.

FIG.16schematically depicts a gearbox1300that can be used with the engines disclosed herein (e.g., the engines100,200,300, and400). The gearbox1300comprises a two-stage star configuration.

The first stage of the gearbox1300includes a first-stage sun gear1302, a first-stage carrier1304housing a plurality of first-stage star gears, and a first-stage ring gear1306. The first-stage sun gear1302can be coupled to a low-speed shaft1308, which in turn is coupled to a low-pressure turbine. The first-stage sun gear1302can mesh with the plurality of first-stage star gears, which mesh with the first-stage ring gear1306. The first-stage carrier1304can be fixed from rotation by a support member1310.

The second stage of the gearbox1300includes a second-stage sun gear1312, a second-stage carrier1314housing a plurality of second-stage star gears, and a second-stage ring gear1316. The second-stage sun gear1312can be coupled to a shaft1318which in turn is coupled to the first-stage ring gear1306. The second-stage carrier1314can be fixed from rotation by a support member1320. The second-stage ring gear1316can be coupled to a fan shaft1322.

In some examples, each stage of the gearbox1300can comprise five star gears. In other examples, the gearbox1300can comprise fewer or more than five star gears in each stage. In some examples, the first-stage carrier1304can comprise a different number of star gears than the second-stage carrier1314. For example, the first-stage carrier1304can comprise five star gears, and the second-stage carrier1314can comprise three star gears, or vice versa.

FIG.17schematically depicts a gearbox1400that can be used with the engines disclosed herein (e.g., the engines100,200,300, and400). The gearbox1400comprises a single-stage star configuration. The gearbox1400includes a sun gear1402, a carrier1404housing a plurality of star gears (e.g., 3-5 star gears), and a ring gear1406. The sun gear1402can mesh with the plurality of star gears, and the plurality of star gears can mesh with the ring gear1406. The sun gear1402can be coupled to a low-speed shaft1408, which in turn is coupled to the low-pressure turbine. The carrier1404can be fixed from rotation by a support member1410. The ring gear1406can be coupled to a fan shaft1412.

FIG.18schematically depicts a gearbox1500that can be used with the engines disclosed herein (e.g., the engines100,200,300, and400). The gearbox1500comprises a single-stage star configuration. The gearbox1500includes a sun gear1502, a carrier1504housing a plurality of star gears (e.g., 3-5 star gears), and a ring gear1506. The sun gear1502can mesh with the plurality of star gears, and the star gears can mesh with the ring gear1506. The sun gear1502can be coupled to a low-speed shaft1508, which in turn is coupled to the low-pressure turbine. The carrier1504can be fixed from rotation by a support member1510. The ring gear1506can be coupled to a fan shaft1512.

FIG.19depicts a gearbox1600that can be used, for example, with the engines disclosed herein (e.g., the engines100,200,400). The gearbox1600is configured as a compound star gearbox. The gearbox1600comprises a sun gear1602and a star carrier1604, which includes a plurality of compound star gears having one or more first portions1606and one or more second portions1608. The gearbox1600further comprises a ring gear1610. The sun gear1602can also mesh with the first portions1606of the plurality of compound star gears. The star carrier can be fixed from rotation via a support member1614. The second portions1608of the plurality of compound star gears can mesh with the ring gear1610. The sun gear1602can be coupled to a low-pressure turbine via the turbine shaft1612. The ring gear1610can be coupled to a fan shaft1616.

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.

Configurations 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 examples, the gear assemblies depicted and described in regard toFIGS.16-19allow 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 configurations of the gear assembly provided herein may allow for gear ratios of up to 10:1. Still various examples of the gear assemblies provided herein may allow for gear ratios within a range of 2.5-4.0. Still yet various examples of the gear assemblies provided herein allow for gear ratios within a range of 4.1-10.0. Other examples can have a gear ratio within a range of 3.0-4.0.FIGS.9,12, and15also provide the gear ratio of several exemplary engines.

Various exemplary gear assemblies are shown and described herein, which can also be referred to as a gearbox. 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 turbomachinery 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 shaft) of the turbomachinery 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 examples, the gear assemblies described herein can be mounted at an aft location (e.g., aft of the combustor and/or the low-pressure turbine).

Portions of a lubricant system1700are depicted schematically inFIG.20. The lubrication system1700can be a component of the turbomachinery engines disclosed herein (e.g., the engines100,200,300, and400) and/or can be coupled to the various gearboxes disclosed herein. For example,FIG.1schematically illustrates the lubricant system coupled to the turbofan engine100and the gear assembly102.FIG.20illustrates a series of lubricant conduits1703can interconnect multiple elements of the lubricant system1700and/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 system1700as shown is by way of example only to illustrate an exemplary system for a turbomachinery engine for circulating lubricant for purposes such as lubrication or heat transfer. Any organization for the lubricant system1700is contemplated, with or without the elements as shown, and/or including additional elements interconnected by any necessary conduit system.

Referring still toFIG.20, the lubricant system1700includes a lubricant reservoir1702configured to store a coolant or lubricant, including organic or mineral oils, synthetic oils, or fuel, or mixtures or combinations thereof. A supply line1704and a scavenge line1706are fluidly coupled to the reservoir1702and collectively form a lubricant circuit to which the reservoir1702and component1710(e.g., a gearbox) can be fluidly coupled. The component1710can be supplied with lubrication by way of a fluid coupling with the supply line1704and can return the supplied lubricant to the reservoir1702by fluidly coupling to the scavenge line1706. More specifically, a component supply line1711can be fluidly coupled between the supply line1704and the component1710. 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 system1700.

Optionally, at least one heat exchanger1705can be included in the lubricant system1700. The heat exchanger1705can 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 turbomachinery engine, fluidly separate from the lubricant passing along the lubricant system1700. 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 system1700utilizing the airflow passing through an outlet guide vane assembly of the turbomachinery engine.

A pump1708can be provided in the lubricant system1700to aid in recirculating lubricant from the reservoir1702to the component1710via the supply line1704. For example, the pump1708can be driven by a rotating component of the turbomachinery engine, such as a high-pressure shaft or a low-pressure shaft of a turbomachinery engine.

Lubricant can be recovered from the component1710by way of the scavenge line1706and returned to the reservoir1702. In the illustrated example, the pump1708is illustrated along the supply line1704downstream of the reservoir1702. The pump1708can be located in any suitable position within the lubricant system1700, including along the scavenge line1706upstream of the reservoir1702. In addition, while not shown, multiple pumps can be provided in the lubricant system1700.

In some examples, a bypass line1712can be fluidly coupled to the supply line1704and scavenge line1706in a manner that bypasses the component1710. In such examples, a bypass valve1715is fluidly coupled to the supply line1704, component supply line1711, and bypass line1712. The bypass valve1715is configured to control a flow of lubricant through at least one of the component supply line1711or the bypass line1712. The bypass valve1715can 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 examples, a plurality of bypass valves can be provided.

During operation, a supply flow1720can move from the reservoir1702, through the supply line1704, and to the bypass valve1715. A component input flow1722can move from the bypass valve1715through the component supply line1711to an inlet of the component1710. A scavenge flow1724can move lubricant from an outlet of the component1710through the scavenge line1706and back to the reservoir1702. Optionally, a bypass flow1726can move from the bypass valve1715through the bypass line1712and to the scavenge line1706. The bypass flow1726can mix with the scavenge flow1724and define a return flow1728moving toward the lubricant reservoir1702.

In one example where no bypass flow exists, it is contemplated that the supply flow1720can be the same as the component input flow1722and that the scavenge flow1724can be the same as the return flow1728. In another example where the bypass flow1726has a nonzero flow rate, the supply flow1720can be divided at the bypass valve1715into the component input flow1722and bypass flow1726. It will also be understood that additional components, valves, sensors, or conduit lines can be provided in the lubricant system1700, and that the example shown inFIG.20is simplified with a single component1710for purposes of illustration.

The lubricant system1700can 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 position1716is located in the supply line1704upstream of the component1710, and a second sensing position1718is located in the scavenge line1706downstream of the component1710.

In one example, the bypass valve1715can 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 valve1715to the first and second sensing positions1716,1718can provide for bypass valve1715sensing or detecting the lubricant temperature at the sensing positions1716,1718as lubricant flows to or from the bypass valve1715. The bypass valve1715can be configured to control the component input flow1722or the bypass flow1726based on the sensed or detected temperature.

It is contemplated that the bypass valve1715, supply line1704, and bypass line1712can at least partially define a closed-loop control system for the component1710. 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 system1700, the bypass valve1715(e.g., mechanical or electrical component) can sense a parameter, such as a lubricant parameter (e.g., temperature), and automatically adjust a system variable, e.g., flow rate to either or both of the bypass line1712or component1710, 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.21, a portion of the lubricant system1700is illustrated supplying lubricant to a particular component1710in the form of a gearbox1750within a turbomachinery engine. The gearbox can be any of the gearboxes disclosed herein. The gearbox1750can include an input shaft1752, an output shaft1754, and a gear assembly1755. In one example, the gear assembly1755can 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 housing1756can at least partially surround the gear assembly1755and form a structural support for the gears and bearings therein. Either or both of the input and output shafts1752,1754can be coupled to the turbomachinery engine. In one example, the input and output shafts1752,1754can 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 line1704can be fluidly coupled to the gearbox1750, such as to the gear assembly1755, to supply lubricant to gears or bearings to the gearbox1750during operation. The scavenge line1706can be fluidly coupled to the gearbox1750, such as to the gear assembly1755or outer housing1756, to collect lubricant. The bypass line1712can be fluidly coupled to the bypass valve1715, supply line1704, and scavenge line1706as shown. A return line1714can also be fluidly coupled to the bypass valve1715, such as for directing the return flow1728to the lubricant reservoir1702for recirculation. While not shown inFIG.21for brevity, the lubricant reservoir1702, the heat exchanger1705, and/or the pump1708(FIG.20) can also be fluidly coupled to the gearbox1750. In this manner, the supply line1704, bypass line1712, scavenge line1706, and return line1714can at least partially define a recirculation line1730(FIG.20) for the lubricant system1700.

The supply flow1720divides at the bypass line into the component input flow1722and the bypass flow1726. In the example shown, the bypass valve1715is in the form of a differential thermal valve that is fluidly coupled to the first and second sensing positions1716,1718.

Lubricant flowing proximate the first and second sensing positions1716,1718provides the respective first and second outputs1741,1742indicative of the temperature of the lubricant at those sensing positions1716,1718. It will be understood that the supply line1704is thermally coupled to the bypass line1712and bypass valve1715such that the temperature of the fluid in the supply line1704proximate the first sensing position1716is approximately the same as fluid in the bypass line1712adjacent the bypass valve1715. 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 valve1715can sense the lubricant temperature in the supply line1704and scavenge line1706via the first and second outputs1741,1742. It can be appreciated that the bypass line1712can form a sensing line for the valve1715to sense the lubricant parameter, such as temperature, at the first sensing position1716.

During operation of the turbomachinery engine, the lubricant temperature can increase within the gearbox1750, such as due to heat generation of the gearbox1750, and throughout the lubricant system1700. In one example, if a lubricant temperature exceeds a predetermined threshold temperature at either sensing position1716,1718, the bypass valve1715can automatically increase the component input flow1722, e.g., from the supply line1704to the gearbox1750, by decreasing the bypass flow1726. Such a predetermined threshold temperature can be any suitable operating temperature for the gearbox1750, such as about 300° F. in some examples. Increasing the component input flow1722can provide for cooling of the gearbox1750, thereby reducing the lubricant temperature sensed in the various lines1704,1706,1712,1714as lubricant recirculates through the lubricant system1700.

In another example, if a temperature difference between the sensing positions1716,1718exceeds a predetermined threshold temperature difference, the bypass valve1715can automatically increase the component input flow1722by decreasing the bypass flow1726. Such a predetermined threshold temperature difference can be any suitable operating temperature for the gearbox1750, 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 positions1716,1718is below the predetermined threshold temperature difference, the bypass valve1715can automatically decrease the component input flow1722or increase the bypass flow1726. In this manner the lubricant system1700can provide for the gearbox1750to operate with a constant temperature difference between the supply and scavenge lines1704,1706.

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

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

A turbomachinery engine comprising a fan assembly, a low-pressure turbine, and a gearbox. The fan assembly includes a plurality of fan blades. The low-pressure turbine comprises 3-5 rotating stages. Each rotating stage of the low-pressure turbine comprises an annular exit area defined by a tip radius of a trailing edge of any one blade of the rotating stage and a hub radius of the any one blade of the rotating stage at an axial location aligned with the tip radius. The low-pressure turbine comprises an area ratio equal to the annular exit area of an aft-most rotating stage of the low-pressure turbine divided by the annular exit area of a forward-most rotating stage of the low-pressure turbine, and the area ratio is within a range of 2.0-6.5. The gearbox includes an input and an output. The input of the gearbox is coupled to the low-pressure turbine and comprises a first rotational speed, and the output of the gearbox is coupled to the fan assembly and comprises a second rotational speed.

The turbomachinery engine of the preceding clause, wherein the area ratio of the low-pressure turbine is within a range of 2.0-3.2.

The turbomachinery engine of any preceding clause, wherein the area ratio of the low-pressure turbine is within a range of 2.2-4.6.

The turbomachinery engine of any preceding clause, wherein the low-pressure turbine includes exactly three rotating stages, and wherein the area ratio of the low-pressure turbine is within a range of 2.2-2.91.

The turbomachinery engine of any preceding clause, wherein the low-pressure turbine includes exactly four rotating stages, and wherein the area ratio of the low-pressure turbine is within a range of 3.1-5.1.

The turbomachinery engine of any preceding clause, wherein the low-pressure turbine includes exactly five rotating stages, and wherein the area ratio of the low-pressure turbine is within a range of 5.4-6.5.

The turbomachinery engine of any preceding clause, wherein the fan assembly is a ducted fan assembly disposed radially within a fan case.

The turbomachinery engine of any preceding clause, wherein the fan assembly is an unducted fan assembly.

The turbomachinery engine of any preceding clause, wherein the fan assembly comprises a first fan and a second fan, each comprising a plurality of fan blades, wherein the second fan is disposed aft of the first fan and has a smaller diameter than the first fan, and wherein the turbomachinery engine is a three-stream engine.

The turbomachinery engine of any preceding clause, wherein the low-pressure turbine comprises an AN2value within a range of 70-104, where A is the annular exit area of the aft-most rotating stage of the low-pressure turbine measured in square inches, N is the rotational speed of the low-pressure turbine measured in revolutions per minute at a redline operating condition, and the product of AN2is divided by 109.

The turbomachinery of any preceding clause, wherein the low-pressure turbine further comprises an area-EGT ratio within a range of 1.3-1.6, wherein the area-EGT

ratio=(the⁢area⁢ratio)(1/(LPT⁢stages-1))(EGT/1000),
where the LPT stages is the number of rotating stages of the low-pressure turbine, and the EGT is an exhaust gas temperature of the low-pressure turbine measured in degrees Celsius at an inlet of the low-pressure turbine at a redline operating condition.

The turbomachinery engine of any preceding clause, wherein an/the exhaust gas temperature of the low-pressure turbine is within a range of 1060-1180 degrees Celsius measured at an/the inlet of the low-pressure turbine at a/the redline operating condition.

The turbomachinery engine of any preceding clause, wherein a gear ratio of the first rotational speed to the second rotational speed is within a range of 2.0-10.0.

The turbomachinery engine of any preceding clause, wherein a gear ratio of the first rotational speed to the second rotational speed is within a range of 2.5-10.0.

The turbomachinery engine of any preceding clause, wherein a gear ratio of the first rotational speed to the second rotational speed is within a range of 2.75-10.0.

The turbomachinery engine of any preceding clause, wherein a gear ratio of the first rotational speed to the second rotational speed is within a range of 2.0-6.0.

The turbomachinery engine of any preceding clause, wherein a gear ratio of the first rotational speed to the second rotational speed is within a range of 3.0-3.5.

Example 18. The turbomachinery engine of any preceding clause, wherein a gear ratio of the first rotational speed to the second rotational speed is within a range of 3.0-4.0.

The turbomachinery engine of any preceding clause, wherein a gear ratio of the first rotational speed to the second rotational speed is within a range of 6.0-10.0.

The turbomachinery engine of any preceding clause, wherein a gear ratio of the first rotational speed to the second rotational speed is within a range of 6.5-9.0.

A turbomachinery engine comprising a fan assembly, a low-pressure turbine, and a gearbox. The fan assembly includes a plurality of fan blades. The low-pressure turbine comprising 3-4 rotating stages. Each rotating stage of the low-pressure turbine comprises an annular exit area defined by a tip radius of a trailing edge of any one blade of the rotating stage and a hub radius of the any one blade of the rotating stage at an axial location aligned with the tip radius. The low-pressure turbine comprises an area ratio equal to the annular exit area of an aft-most rotating stage of the low-pressure turbine divided by the annular exit area of a forward-most rotating stage of the low-pressure turbine, and the area ratio is within a range of 2.1-4.6. The gearbox including an input and an output. The input of the gearbox is coupled to the low-pressure turbine and comprises a first rotational speed, the output of the gearbox 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 3.0-3.5.

The turbomachinery engine of any preceding clause, wherein the area ratio of the low-pressure turbine is within a range of 2.2-3.2.

The turbomachinery engine of any preceding clause, wherein the low-pressure turbine includes exactly three rotating stages, and wherein the area ratio of the low-pressure turbine is within a range of 2.2-2.99.

The turbomachinery engine of any preceding clause, wherein the low-pressure turbine includes exactly four rotating stages, and wherein the area ratio of the low-pressure turbine is within a range of 2.3-4.59.

The turbomachinery engine of any preceding clause, wherein the fan assembly is a ducted fan assembly.

The turbomachinery engine of any preceding clause, wherein the ducted fan assembly comprises a first ducted fan and a second ducted fan, each comprising a plurality of fan blades, wherein the second ducted fan is disposed aft of the first ducted fan and has a smaller diameter than the first ducted fan, and wherein the turbomachinery engine is a three-stream engine.

The turbomachinery engine of any preceding clause, wherein the low-pressure turbine comprises an AN2value within a range of 70-104, where A the annular exit area of the aft-most rotating stage of the low-pressure turbine measured in square inches, N is the rotational speed of the low-pressure turbine measured in revolutions per minute at a redline operating condition, and a product of AN2is divided by 109.

The turbomachinery of any preceding clause, wherein the low-pressure turbine further comprises an area-EGT ratio within a range of 1.3-1.6, wherein the area-EGT

ratio=(the⁢area⁢ratio)(1/(LPT⁢stages-1))(EGT/1000),
where the LPT stages is a number of rotating stages of the low-pressure turbine, and the EGT is an exhaust gas temperature of the low-pressure turbine measured in degrees Celsius at an inlet of the low-pressure turbine at a redline operating condition.

The turbomachinery engine of any preceding clause, wherein an/the exhaust gas temperature of the low-pressure turbine is within a range of 1060-1180 degrees Celsius measured at an/the inlet of the low-pressure turbine at a/the redline operating condition.

The turbomachinery engine of any preceding clause, wherein the fan assembly comprises 8-22 fan blades, and wherein the turbomachinery engine further comprises a low-pressure compressor comprising 1-8 stages, a high-pressure compressor comprising 8-11 stages, and a high-pressure turbine comprising 1-2 stages.

The turbomachinery engine of any preceding clause, wherein: the fan assembly comprises 12-18 fan blades; the low-pressure compressor comprises 3-5 stages; the high-pressure compressor comprises 8-9 stages; and the high-pressure turbine comprises 2 stages.

A turbomachinery engine comprising a fan assembly, a low-pressure turbine, and a gearbox. The fan assembly includes a plurality of fan blades. The low-pressure turbine comprises 4-5 rotating stages. Each rotating stage of the low-pressure turbine comprises an annular exit area defined by a tip radius of a trailing edge of any one blade of the rotating stage and a hub radius of the any one blade of the rotating stage at an axial location aligned with the tip radius. The low-pressure turbine comprises an area ratio equal to the annular exit area of an aft-most rotating stage of the low-pressure turbine divided by the annular exit area of a forward-most rotating stage of the low-pressure turbine, and the area ratio is within a range of 3.6-5.79. The gearbox includes an input and an output. The input of the gearbox is coupled to the low-pressure turbine and comprises a first rotational speed, the output of the gearbox 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 5.0-10.0.

The turbomachinery engine of any preceding clause, wherein the gear ratio is within a range of 6.0-9.0.

The turbomachinery engine of any preceding clause, wherein the area ratio of the low-pressure turbine is within a range of 4.1-5.79.

The turbomachinery engine of any preceding clause, wherein the low-pressure turbine includes exactly four rotating stages.

The turbomachinery engine of any example herein, and particularly any one of examples 32-34, wherein the low-pressure turbine includes exactly five rotating stages.

The turbomachinery engine of any preceding clause, wherein the fan assembly is an unducted fan assembly.

The turbomachinery engine of any preceding clause, further comprising a ducted fan assembly disposed aft of the unducted fan assembly, and wherein the turbomachinery engine is a three-stream engine.

The turbomachinery engine of any preceding clause, wherein the low-pressure turbine comprises an AN2value within a range of 70-104, where A is the annular exit area of the aft-most rotating stage of the low-pressure turbine measured in square inches, N is the rotational speed of the low-pressure turbine measured in revolutions per minute at a redline operating condition, and a product of AN2is divided by 109.

The turbomachinery of any preceding clause, wherein the low-pressure turbine further comprises an area-EGT ratio within a range of 1.3-1.6, wherein the area-EGT

ratio=(the⁢area⁢ratio)(1/(LPT⁢stages-1))(EGT/1000),
where the LPT stages is a number of rotating stages of the low-pressure turbine, and the EGT is an exhaust gas temperature of the low-pressure turbine measured in degrees Celsius at an inlet of the low-pressure turbine at a redline operating condition.

The turbomachinery engine of any preceding clause, wherein an/the exhaust gas temperature of the low-pressure turbine is within a range of 1060-1180 degrees Celsius measured at an/the inlet of the low-pressure turbine at a/the redline operating condition.

The turbomachinery engine of any preceding clause, wherein the fan assembly comprises 8-22 fan blades. The turbomachinery engine further comprises a low-pressure compressor comprising 1-5 stages, a high-pressure compressor comprising 7-11 stages, a high-pressure turbine comprising 1-2 stages.

The turbomachinery engine of any preceding clause, wherein: the fan assembly comprises 12-18 fan blades; the low-pressure compressor comprises 3-5 stages; the high-pressure compressor comprises 8-10 stages; and the high-pressure turbine comprises 2 stages.

A turbomachinery engine comprising a fan assembly, a low-pressure turbine, and a gearbox. The fan assembly including a plurality of fan blades. The low-pressure turbine comprises 3-5 rotating stages and an area-EGT ratio within a range of 1.3-1.6. The area-EGT

ratio=(area⁢ratio)(1/(LPT⁢stages-1))(EGT/1000).
Each rotating stage of the low-pressure turbine comprises an annular exit area defined by a tip radius of a trailing edge of any one blade of the rotating stage and a hub radius of the any one blade of the rotating stage at an axial location aligned with the tip radius. The area ratio is the annular exit area of an aft-most rotating stage of the low-pressure turbine divided by the annular exit area of a forward-most rotating stage of the low-pressure turbine. The LPT stages is the number of rotating stages of the low-pressure turbine. The EGT is an exhaust gas temperature of the low-pressure turbine measured in degrees Celsius at an inlet of the low-pressure turbine at a redline operating condition. The gearbox including an input and an output. The input of the gearbox is coupled to the low-pressure turbine and comprises a first rotational speed, and the output of the gearbox is coupled to the fan assembly and has a second rotational speed.

The turbomachinery engine of any preceding clause, wherein the area-EGT ratio is within a range of 1.38-1.58.

The turbomachinery engine of any preceding clause, wherein the area-EGT ratio is within a range of 1.31-1.53.

The turbomachinery engine of any preceding clause, wherein the area-EGT ratio is within a range of 1.30-1.36.

The turbomachinery engine of any preceding clause, wherein the area ratio of the low-pressure turbine is within a range of 2.2-4.6.

The turbomachinery engine of any preceding clause, wherein the low-pressure turbine includes exactly three rotating stages, and wherein the area ratio of the low-pressure turbine is within a range of 2.2-2.91.

The turbomachinery engine of any preceding clause, wherein the low-pressure turbine includes exactly four rotating stages, and wherein the area ratio of the low-pressure turbine is within a range of 3.0-5.2.

The turbomachinery engine of any preceding clause, wherein the low-pressure turbine includes exactly five rotating stages, and wherein the area ratio of the low-pressure turbine is within a range of 5.0-6.5.

The turbomachinery engine of any preceding clause, wherein the fan assembly is a ducted fan assembly disposed radially within a fan case.

The turbomachinery engine of any preceding clause, wherein the fan assembly is an unducted fan assembly.

The turbomachinery engine of any preceding clause, wherein the fan assembly comprises a first fan and a second fan, each comprising a plurality of fan blades, wherein the second fan is disposed aft of the first fan and has a smaller diameter than the first fan, and wherein the turbomachinery engine is a three-stream engine.

The turbomachinery engine of any preceding clause, wherein the low-pressure turbine comprises an AN2value within a range of 70-104, where A is the annular exit area of the aft-most rotating stage of the low-pressure turbine measured in square inches, N is the rotational speed of the low-pressure turbine measured in revolutions per minute at a redline operating condition, and a product of AN2is divided by 109.

The turbomachinery engine of any preceding clause, wherein the exhaust gas temperature of the low-pressure turbine is within a range of 1060-1180 degrees Celsius measured at the inlet of the low-pressure turbine at the redline operating condition.

The turbomachinery engine of any preceding clause, wherein a gear ratio of the first rotational speed to the second rotational speed is within a range of 2.0-10.0.

The turbomachinery engine of any preceding clause, wherein a gear ratio of the first rotational speed to the second rotational speed is within a range of 2.5-10.0.

The turbomachinery engine of any preceding clause, wherein a gear ratio of the first rotational speed to the second rotational speed is within a range of 2.75-10.0.

The turbomachinery engine of any preceding clause, wherein a gear ratio of the first rotational speed to the second rotational speed is within a range of 2.75-3.5.

The turbomachinery engine of any preceding clause, wherein a gear ratio of the first rotational speed to the second rotational speed is within a range of 3.0-4.0.

The turbomachinery engine of any preceding clause, wherein a gear ratio of the first rotational speed to the second rotational speed is within a range of 4.1-10.0.

The turbomachinery engine of any preceding clause, wherein a gear ratio of the first rotational speed to the second rotational speed is within a range of 6.0-10.0.

The turbomachinery engine of any preceding clause, wherein a gear ratio of the first rotational speed to the second rotational speed is within a range of 6.5-9.0.

A turbine for an aircraft engine comprising 3-5 rotating stages and an area ratio within a range of 2.0-6.5. Each rotating stage comprises an annular exit area defined by a tip radius of a trailing edge of any one blade of the rotating stage and a hub radius of the any one blade of the rotating stage at an axial location aligned with the tip radius. The area ratio equals the annular exit area of an aft-most rotating stage divided by the annular exit area of a forward-most rotating stage.

The turbine of any preceding clause, wherein the turbine is a low-pressure turbine disposed aft of a high-pressure turbine.

A turbine for an aircraft engine comprising 3-5 rotating stages and an area-EGT ratio within a range of 1.3-1.6. The area-EGT

ratio=(area⁢ratio)(1/(stages-1))(EGT/1000).
Each rotating stage comprises an annular exit area defined by a tip radius of a trailing edge of any one blade of the rotating stage and a hub radius of the any one blade of the rotating stage at an axial location aligned with the tip radius. The area ratio is the annular exit area of an aft-most rotating stage divided by the annular exit area of a forward-most rotating stage, wherein the stages is the number of rotating stages. The EGT is an exhaust gas temperature measured in degrees Celsius at an inlet of the turbine at a redline operating condition.

The turbine of any preceding clause, wherein the turbine is a low-pressure turbine disposed aft of a high-pressure turbine.