Open rotor turbomachinery engine

Abstract

A turbomachinery engine comprising a non-ducted rotor assembly, a low-pressure turbine, a low-pressure shaft, a gearbox, and engine-related parameters. The non-ducted rotor assembly comprises a plurality of rotor blades arranged in a single row and defining a blade diameter. The gearbox comprises an input, an output, and a gear ratio. The input of the gearbox is coupled to the low-pressure shaft (which is coupled to the LPT), and the output of the gearbox is coupled to the non-ducted rotor assembly. The engine-related parameters are greater than 0.17 and less than 0.83. The engine-related parameters are equal to D / N / GR, where D is the blade tip diameter measured in feet, N is the number of stages of the low-pressure turbine, and GR is the gear ratio of the gearbox. The non-ducted propeller assembly optionally includes 60‑180 HP / ft under takeoff flight conditions. 2 range of disc loads.

Description

Technical Field

[0001] The present disclosure relates generally to turbomachinery engines, and in particular, to open rotor turbomachinery engine configurations. Background Art

[0002] A turbomachinery engine used in an aircraft consists of a core engine that drives a propeller. The propeller generates most of the turbomachinery engine's thrust. The thrust generated by the propeller can be used to move the aircraft.

[0003] In some cases, a turbomachinery engine is configured as a direct drive engine. A direct drive engine is configured such that the power turbine (e.g., low-pressure turbine) of the core engine is directly coupled to the propeller. As a result, the power turbine and the propeller rotate at the same speed (i.e., the same rpm).

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

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

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

[0007] In certain cases, a turbomachinery engine may include a gearbox and an exposed propeller. Despite having certain advantages, a geared open rotor engine may have one or more disadvantages; for example, including a gearbox in a turbomachinery engine may introduce additional complexity to the engine. For example, this may make engine development and / or manufacturing more difficult. Additionally or alternatively, an open rotor engine may generate undesirable noise levels. Therefore, there is a need for improved geared open rotor turbomachinery engines. There is also a need for apparatus and methods that can be used to more efficiently and / or more accurately develop and manufacture geared open rotor turbomachinery engines. Summary of the Invention

[0008] Aspects and advantages of the disclosed technology will be set forth in part in the description which follows, or may be obvious from the description, or may be learned through practice of the technology disclosed in the description.

[0009] Various turbomachinery engines and gear assemblies are disclosed herein. The disclosed turbomachinery engines include a gearbox and an open rotor propeller. In addition, the disclosed turbomachinery engines are characterized or defined by engine-related parameters. The engine-related parameter (ECP) is equal to D / N / GR, where D is the blade tip diameter measured in feet, where N is the number of stages of the low-pressure turbine, and where GR is the gear ratio of the gearbox. For example, ECP can be used to construct an engine so that the engine has higher efficiency and lower noise than a typical engine. In some cases, the engines disclosed herein including ECP are lighter, have higher propulsion efficiency, have higher engine efficiency and / or are quieter than a typical engine. ECP therefore provides an improved turbomachinery engine and / or can help simplify one or more complexities of geared open rotor engine development.

[0010] In certain embodiments, a turbomachinery engine includes a non-ducted rotor assembly (which may also be referred to as a non-ducted propeller assembly), a low-pressure turbine, a low-pressure shaft, a gearbox, and engine-related parameters. The non-ducted rotor assembly includes a plurality of rotor blades arranged in a single row and defining a blade diameter. The low-pressure turbine includes a number of stages defined by the rows of rotating blades of the low-pressure turbine. The low-pressure shaft is coupled to the low-pressure turbine. The gearbox includes an input, an output, and a gear ratio. The input of the gearbox is coupled to the low-pressure shaft, and the output of the gearbox is coupled to the non-ducted rotor assembly. The engine-related parameters are greater than 0.17 and less than 0.83.

[0011] In some examples, a single non-ducted rotor turbomachinery engine includes a low-pressure turbine, a low-pressure spool, a gearbox, and a non-ducted rotor assembly. The low-pressure turbine defines a number of stages. The low-pressure spool is coupled to the low-pressure turbine. The gearbox defines a gear ratio. The non-ducted rotor assembly includes a plurality of rotor blades defining a blade tip diameter, and the non-ducted rotor assembly is coupled to the low-pressure spool via the gearbox. The single non-ducted rotor turbomachinery engine defines an engine-related parameter equal to the blade tip diameter measured in feet divided by the number of low-pressure turbine stages divided by the gear ratio of the gearbox, and the engine-related parameter is greater than 0.17 and less than 0.83.

[0012] In another example, a turbomachinery engine includes a non-ducted propeller assembly, a low-pressure compressor, a low-pressure turbine, a low-pressure shaft, a high-pressure compressor, a high-pressure turbine, a high-pressure shaft, a gearbox, and engine-related parameters. The non-ducted propeller assembly includes a plurality of propeller blades arranged in a single row and defining a blade diameter, and the blade tip diameter is in the range of 8-14 feet or 10-16 feet. The low-pressure compressor includes one or more LPC rotors defining a number of LPC stages. The low-pressure turbine includes multiple LPT rotors defining a number of LPT stages, and the number of LPT stages is in the range of 3-8. The low-pressure shaft is coupled to the low-pressure compressor and the low-pressure turbine. The high-pressure compressor includes multiple HPC rotors defining a number of HPC stages. The high-pressure turbine includes one or more HPT rotors defining a number of HPT stages. The high-pressure shaft is coupled to the high-pressure compressor and the high-pressure turbine. The gearbox includes an input, an output, and a gear ratio. The input is coupled to the low-pressure shaft and is configured to rotate at a first speed. The output is coupled to the non-ducted propeller assembly and is configured to rotate at a second speed less than the first speed. The gear ratio is defined by a ratio of the first speed to the second speed and is in the range of 4 to 12. The engine-related parameter is in the range of 0.17 to 0.83, where the engine-related parameter is equal to D / N / GR, where D is the blade tip diameter of the propeller blade measured in feet, N is the number of LPT stages of the low pressure turbine, and GR is the gear ratio of the gearbox.

[0013] The propulsor of the non-ducted engine disclosed herein may also include a specific range of disc loads under takeoff flight conditions. More specifically, the disclosed engine may include a propulsor having a range of 60-180 HP / ft2 under takeoff flight conditions. 2 Disc loading is the power measured in horsepower (HP) divided by the square feet (ft) of air under takeoff flight conditions. 2 ) is the swept area of ​​the propeller measured in units of ). In some cases, specific propeller disc load ranges and ECP ranges can be combined to provide improved engine configurations. For example, in some embodiments, an engine may include an ECP of 0.17-0.83 and 60-180 HP / ft at takeoff flight conditions. 2 In other examples, the engine may include an ECP of 0.17-0.63 and a propeller disc load of 60-180 HP / ft at takeoff flight conditions. 2 An engine including the disclosed ECP and / or propeller disc load ranges may, for example, advantageously provide improved engine performance and efficiency compared to typical engines while also meeting acoustic requirements.

[0014] These and other features, aspects and / or advantages of the present disclosure will become better understood with reference to the following description, drawings and claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the disclosed technology and, together with the description, serve to explain the principles of the present disclosure. BRIEF DESCRIPTION OF THE DRAWINGS

[0015] Figure 1 is a schematic cross-sectional view of an exemplary embodiment of a turbomachinery engine configured with an open rotor propulsion system.

[0016] Figure 2 is a cross-sectional schematic diagram of an exemplary embodiment of a turbomachinery engine including an open rotor propulsion system, a three-flow architecture, and one or more heat exchangers in a third flow of the three-flow architecture.

[0017] Figure 3 is a schematic cross-sectional view of an exemplary embodiment of a counter-rotating low-pressure turbine of a turbomachinery engine, the low-pressure turbine having a 3x3 configuration.

[0018] Figure 4 is a schematic cross-sectional view of an exemplary embodiment of a counter-rotating low-pressure turbine of a turbomachinery engine, the low-pressure turbine having a 4x3 configuration.

[0019] Figure 5 is a schematic cross-sectional view of an exemplary embodiment of a gearbox configuration for a turbomachinery engine.

[0020] Figure 6 is a schematic cross-sectional view of an exemplary embodiment of a gearbox configuration for a turbomachinery engine.

[0021] Figure 7 is a schematic cross-sectional view of an exemplary embodiment of a gearbox configuration for a turbomachinery engine.

[0022] Figure 8 is a schematic cross-sectional view of an exemplary embodiment of a gearbox configuration for a turbomachinery engine.

[0023] Figure 9 is a schematic cross-sectional view of an exemplary embodiment of a gearbox configuration for a turbomachinery engine.

[0024] Figure 10 is a graph depicting various engine parameters for several exemplary turbomachinery engines including a 10 foot fan diameter.

[0025] Figure 11 is a graph depicting various engine parameters for several exemplary turbomachinery engines including an 11 foot fan diameter.

[0026] Figure 12 is a graph depicting various engine parameters for several exemplary turbomachinery engines including a 12 foot fan diameter.

[0027] Figure 13 is a graph depicting various engine parameters for several exemplary turbomachinery engines including a 13 foot fan diameter.

[0028] Figure 14 is a graph depicting various engine parameters for several exemplary turbomachinery engines including a 14 foot fan diameter.

[0029] Figure 15 is a graph depicting various engine parameters for several exemplary turbomachinery engines including a 15 foot fan diameter.

[0030] Figure 16 is a graph depicting various engine parameters for several exemplary turbomachinery engines including a 16 foot fan diameter.

[0031] Figure 17 is a graph depicting exemplary ranges of engine-related parameters relative to exemplary ranges of propeller disk loading for a turbomachinery engine.

[0032] Figure 18 is a graph depicting exemplary ranges of engine-related parameters relative to exemplary ranges of propeller disk loading for a turbomachinery engine. DETAILED DESCRIPTION

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

[0034] The word “exemplary” is used herein to mean “serving as an example, instance, or illustration.” Any implementation described herein as “exemplary” is not necessarily to be construed as preferred or advantageous over other implementations.

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

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

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

[0038] Unless otherwise indicated herein, the terms "coupled," "fixed," "attached," and the like refer to both direct coupling, fixing, or attachment as well as indirect coupling, fixing, or attachment through one or more intermediate components or features.

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

[0040] As used throughout the specification and claims, approximating language is used to modify any quantitative expression that can be permitted to vary without resulting in a change in the basic function to which it is related. Thus, values ​​modified by terms such as "about," "approximately," and "substantially" are not limited to the precise values ​​specified. In at least some cases, approximate language can correspond to the precision of an instrument used to measure a value, or the precision of a method or machine used to construct or manufacture a component and / or system. For example, approximate language can refer to within a margin of 1%, 2%, 4%, 5%, 10%, 15%, or 20% of a single value, a range of values, and / or an endpoint of a defined range of values.

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

[0042] Referring now to the accompanying drawings, Figure 11 is an exemplary embodiment of an engine 100 including a gear assembly 102 according to aspects of the present disclosure. The engine 100 includes a fan assembly 104 driven by a core engine 106. In various embodiments, the core engine 106 is a Brayton cycle system configured to drive the fan assembly 104. The core engine 106 is at least partially covered by a casing 114. The fan assembly 104 includes a plurality of fan blades 108. The vane assembly 110 extends from the casing 114 in a cantilevered manner. Therefore, the vane assembly 110 may also be referred to as a non-ducted vane assembly. The vane assembly 110, including a plurality of stationary outlet guide vanes 112, is positioned in an operative arrangement with the fan blades 108 to provide thrust, control thrust vectoring, reduce or redirect undesirable noise, and / or otherwise desirably change the flow of air relative to the fan blades 108.

[0043] In some embodiments, fan assembly 104 includes eight (8) to twenty (20) fan blades 108 (which may also be referred to as "propeller blades"). In particular embodiments, fan assembly 104 includes ten (10) to eighteen (18) fan blades 108. In certain embodiments, fan assembly 104 includes twelve (12) to sixteen (16) fan blades 108. In some embodiments, blade assembly 110 includes three (3) to thirty (30) blades 112. In certain embodiments, blade assembly 110 includes the same number of blades 112 as fan blades 108 or a fewer number of blades 112 than fan blades 108. For example, in particular embodiments, engine 100 includes twelve (12) fan blades 108 and ten (10) blades 112. In other embodiments, blade assembly 110 includes a greater number of blades 112 than fan blades 108. For example, in particular embodiments, engine 100 includes ten (10) fan blades 108 and twenty-three (23) blades 112.

[0044] In certain embodiments, as Figure 1 As shown, the blade assembly 110 is positioned downstream or rearward of the fan assembly 104. However, it should be understood that in some embodiments, the blade assembly 110 can be positioned upstream or forward of the fan assembly 104. In various embodiments, the engine 100 may include a first blade assembly positioned in front of the fan assembly 104 and a second blade assembly positioned rearward of the fan assembly 104. The fan assembly 104 can be configured to adjust the pitch of one or more fan blades 108 as desired, for example, to control thrust vectoring, reduce or redirect noise, and / or change thrust output. The blade assembly 110 can be configured to adjust the pitch of one or more blades 112 as desired, for example, to control thrust vectoring, reduce or redirect noise, and / or change thrust output. The pitch control mechanisms at one or both of the fan assembly 104 or the blade assembly 110 can cooperate to produce one or more of the desired effects described above.

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

[0046] Fan blade 108 includes a diameter (D fan ). It should be noted that for illustration purposes only D is shown. fan Half of the fan tip diameter (D fan ) can be within the range of 8-20 feet. For example, in some embodiments, D fan In certain embodiments, D fan In some embodiments, D fan 12-13 feet.

[0047] In some embodiments, the fan blade tip speed at cruise flight conditions can be 500 to 900 feet per second (fps), or 700 to 800 fps. At cruise flight conditions, the fan can rotate at a speed of 400-1200 rpm (or 450-1000 rpm). As measured across the fan blades at cruise flight conditions, the fan pressure ratio (FPR) of fan assembly 104 can be 1.04 to 1.10, or in some embodiments, 1.05 to 1.08.

[0048] The fans disclosed herein may include a robustness. The robustness is based on the average blade chord, which is defined as the blade planar area (the surface area on one side of the blade) divided by the blade radial span. The robustness is proportional to the number of blades and the chord length, and inversely proportional to the diameter. For the purposes of this disclosure, the robustness is equal to the product of the average blade chord (C) times the number of fan blades (N) divided by two (2) times pi (π) times a reference radius (R_ref), which is herein a radius equal to 0.75 times the tip radius (Rt) of the rotor blade (i.e., CxN / (2xπxR_ref)). Using this formula, the fan may include a robustness between 0.5 and 1.0, or more specifically, between 0.6 and 1. In other examples, the fan may include a robustness between 1.1 and 1.5, or in some examples, between 1.1-1.3.

[0049] Cruising altitude is generally the altitude of an aircraft after climbing and before descending to a level before the approach phase of flight. In various embodiments, the engine is applied to a vehicle with a cruising altitude of up to approximately 65,000 ft. In certain embodiments, the cruising altitude is between approximately 28,000 ft and approximately 45,000 ft. In certain embodiments, the cruising speed is in the range of approximately Mach 0.6-0.9 (or alternatively, approximately Mach 0.7-0.9). In certain embodiments, the cruising altitude is expressed in flight levels (FL) based on standard air pressure at sea level, wherein cruising flight conditions are between FL280 and FL650. In another embodiment, cruising flight conditions are between FL280 and FL450. In certain embodiments, the cruising altitude is defined at least based on air pressure, wherein the cruising altitude is between approximately 4.85 psia and approximately 0.82 psia based on a sea level pressure of approximately 14.70 psia and a sea level temperature of approximately 59 degrees Fahrenheit. In another embodiment, the cruising altitude is between approximately 4.85 psia and approximately 2.14 psia. It will be appreciated that in certain embodiments, the cruising altitude range defined by pressure may be adjusted based on different reference sea level pressures and / or sea level temperatures.

[0050] The core engine 106 is generally enclosed in a casing 114 that defines a core diameter (D core ), which can be considered as the maximum extent from the centerline axis (Rbase). In certain embodiments, the engine 100 includes a length (L) from a longitudinal (or axial) front end 116 to a longitudinal rear end 118. In various embodiments, the engine 100 defines an L / D that provides reduced installation drag. core In one embodiment, L / D core At least 2. In another embodiment, L / D core At least 2.5. In some embodiments, L / D coreLess than 5, less than 4 and less than 3. In various embodiments, it should be understood that L / D core For use with single non-ducted rotary engines.

[0051] The reduced installed drag can further provide improved efficiency, such as improved specific fuel consumption. Additionally or alternatively, the reduced drag can provide engine and aircraft operation at a cruising altitude equal to or above Mach 0.5. In some embodiments, L / D core The fan assembly 104 and / or the blade assembly 110, individually or together, at least partially configure the engine 100 to operate at a maximum cruise altitude operating speed between approximately Mach 0.55 and approximately Mach 0.85; or between approximately 0.72 and 0.85, or between approximately 0.75 and 0.85. For example, the engine 100 can be configured to operate at a maximum cruise altitude operating speed between approximately Mach 0.6 and approximately Mach 0.9.

[0052] Still refer to Figure 1 , core engine 106 extends in a radial direction (R) relative to engine centerline axis 120. Gear assembly 102 receives power or torque from core engine 106 via power input source 122 and provides power or torque via power output source 124 to drive fan assembly 104 in a circumferential direction C about engine centerline axis 120.

[0053] The gear assembly 102 of the engine 100 may include a plurality of gears, including an input and an output. The gear assembly may also include one or more intermediate gears disposed between and / or interconnecting the input and output. The input may be coupled to a turbine section of the core engine 106 and may have a first rotational speed. The output may be coupled to a fan assembly and may have a second rotational speed. In some embodiments, the gear ratio of the first rotational speed to the second rotational speed is greater than 4.1 (e.g., within a range of 4.1-14.0).

[0054] The gear assembly 102 (which may also be referred to as a "gearbox") may include various types and / or configurations. For example, in some embodiments, the gearbox is an epicyclic gearbox configured as a planetary gear configuration. The planetary gear configuration includes a sun gear, a plurality of planetary gears (which may also be referred to as "planetary gears"), and a ring gear. The sun gear is the input and is coupled to a power turbine (e.g., a low-pressure turbine) such that the sun gear and the power turbine rotate at the same speed. The planetary gears are disposed between the sun gear and the ring gear and are interconnected. The planetary gears are rotatably coupled to a fixed carrier. Therefore, the planetary gears can rotate about their respective axes but cannot orbit together relative to the sun gear or the ring gear. As another example, the gearbox is an epicyclic gearbox configured as a planetary gear configuration. The planetary gear configuration includes a sun gear, a plurality of planetary gears, and a ring gear. The sun gear is the input and is coupled to the power turbine. The planetary gears are disposed between the sun gear and the ring gear and are interconnected. The planetary gears are rotatably coupled to a rotatable carrier. The planet gears can therefore rotate about their respective axes and can also rotate together with the carrier relative to the sun gear and ring gear. The carrier is the output and is coupled to the fan assembly. The ring gear is fixed and does not rotate.

[0055] In some embodiments, the gearbox is a single-stage gearbox (e.g., Figure 10-11 In other embodiments, the gearbox is a multi-stage gearbox (e.g., Figure 9 and 12 In some embodiments, the gearbox is an epicyclic gearbox. In some embodiments, the gearbox is a non-episodic gearbox (e.g., a compound gearbox- Figure 13 ).

[0056] As described above, gear assembly can be used to reduce the speed of output relative to input. In certain embodiments, the gear ratio of input speed to output speed is greater than 4.1. For example, in a specific embodiment, the gear ratio is in the range of 4.1-14.0, in the range of 4.5-14.0 or in the range of 6.0-14.0. In certain embodiments, the gear ratio is in the range of 4.5-12 or in the range of 6.0-11.0. Therefore, in certain embodiments, fan assembly can be configured to rotate at a speed of 400-1500rpm under cruise flight conditions, and power turbine (e.g., low-pressure turbine) is configured to rotate at a speed of 2,500-15,000rpm under cruise flight conditions. In a specific embodiment, fan assembly can be configured to rotate at a speed of 450-1200rpm under cruise flight conditions, and power turbine is configured to rotate at a speed of 5,000-10,000rpm under cruise flight conditions.

[0057] exist Figure 5-9Various gear assembly configurations are schematically depicted in FIG. These gearboxes may be used with any engine disclosed herein, including engine 100. More detailed information regarding the gearboxes is provided below.

[0058] Figure 2 A cross-sectional view of engine 200 is shown, which is configured as an exemplary embodiment of an open rotor propulsion engine. Engine 200 is generally similar to engine 100 and corresponding components have been numbered similarly. For example, the gear assembly of engine 100 is numbered "102," the gear assembly of engine 200 is numbered "202," and so on. In addition to gear assembly 202, engine 200 includes a fan assembly 204, which includes a plurality of fan blades 208 distributed about an engine centerline axis 220. In some cases, fan blades 208 are circumferentially arranged in an equally spaced relationship about engine centerline axis 220, and each fan blade 208 has a root 225 and a tip 226, an axial span defined therebetween, and a center blade axis 228. In other cases, the fan blades may be non-equally spaced.

[0059] The core engine 206 includes a compressor section 230, a combustion section 232, and a turbine section 234 (which may be referred to as an "expansion section") arranged together in series flow. The core engine 206 extends circumferentially relative to the engine centerline axis 220. The core engine 206 includes a high-speed spool including a high-pressure compressor 236 and a high-pressure turbine 238 operably and rotatably coupled together by a high-speed shaft 240. The combustion section 232 is positioned between the high-pressure compressor 236 and the high-pressure turbine 238.

[0060] The combustion section 232 can be configured as a deflagration combustion section, a rotating detonation combustion section, a pulse detonation combustion section, and / or other suitable heat addition systems. The combustion section 232 can be configured as one or more of a rich burn system or a lean burn system, or a combination thereof. In still various embodiments, the combustion section 232 includes an annular combustor, a can combustor, a tubular combustor, a trapped vortex combustor (TVC), or other suitable combustion systems, or a combination thereof.

[0061] Core engine 206 also includes a supercharger or low-pressure compressor positioned in flow relationship with high-pressure compressor 236. Low-pressure compressor 242 is rotatably coupled to low-pressure turbine 244 via a low-speed shaft 246 to enable low-pressure turbine 244 to drive low-pressure compressor 242. Low-speed shaft 246 is also operatively connected to gear assembly 202 to provide power to fan assembly 204, such as described further herein.

[0062] It should be understood that the terms "low" and "high" or their respective comparatives (e.g., "lower" and "higher," as applicable), when used with compressor, turbine, shaft, or spool components, refer to relative pressures and / or relative speeds within the engine, unless otherwise specified. For example, a "low spool" or "low-speed shaft" defines a component that is configured to operate at a speed (e.g., maximum allowable speed) lower than the "high spool" or "high-speed shaft" of the engine. Alternatively, the above terms may be understood as their superlatives, unless otherwise specified. For example, a "low turbine" or "low-speed turbine" may refer to the lowest maximum speed turbine within the turbine section, a "low compressor" or "low-speed compressor" may refer to the lowest maximum speed turbine within the compressor section, a "high turbine" or "high-speed turbine" may refer to the highest maximum speed turbine within the turbine section, and a "high compressor" or "high-speed compressor" may refer to the highest maximum speed compressor within the compressor section. Similarly, a low-speed spool refers to a lower maximum speed than a high-speed spool. It should also be understood that the terms "low" or "high" in the above aspects may additionally or alternatively be understood as relative to a minimum allowable speed, or relative to a minimum or maximum allowable speed for normal, desired, steady-state engine operation, etc.

[0063] The compressor and / or turbine disclosed herein may include various numbers of stages. As disclosed herein, the number of stages includes the number of rotor or blade stages in a particular component (e.g., a compressor or turbine). For example, in some embodiments, the low-pressure compressor may include 1-8 stages, the high-pressure compressor may include 8-15 stages, the high-pressure turbine may include 1-2 stages, and / or the low-pressure turbine may include 3-8 stages (or alternatively 4-8 LPT stages, or alternatively 4-6 LPT stages). For example, in certain embodiments, the engine may include a one-stage low-pressure compressor, an 11-stage high-pressure compressor, a two-stage high-pressure compressor, and a seven-stage low-pressure turbine. As another example, the engine may include a three-stage low-pressure compressor, a 10-stage high-pressure compressor, a two-stage high-pressure compressor, and a seven-stage low-pressure turbine. As yet another example, the engine may include a three-stage low-pressure compressor, a 10-stage high-pressure compressor, a two-stage high-pressure compressor, and a four-stage low-pressure turbine.

[0064] In some embodiments, the low-pressure turbine is a counter-rotating low-pressure turbine comprising inner bladed stages and outer bladed stages. The inner bladed stages extend radially outward from the inner shaft, and the outer bladed stages extend radially inward from the outer drum. In certain embodiments, the counter-rotating low-pressure turbine comprises three inner bladed stages and three outer bladed stages, which may be collectively referred to as a six-stage low-pressure turbine. In other embodiments, the counter-rotating low-pressure turbine comprises four inner bladed stages and three outer bladed stages, which may be collectively referred to as a seven-stage low-pressure turbine.

[0065] As discussed in more detail below, core engine 206 includes a gear assembly 202 configured to transfer power from turbine section 234 and reduce the output rotational speed at fan assembly 204 relative to low-pressure turbine 244. The embodiments of gear assembly 202 depicted and described herein can allow for gear ratios suitable for large diameter non-ducted fans (e.g., 4.1-14.0, 4.5-14.0, and / or 6.0-14.0 gear ratios). Furthermore, the embodiments of gear assembly 202 provided herein can be adapted to the radial or diametrical constraints of core engine 206 within casing 214.

[0066] exist Figure 5-9 Various gearbox configurations are schematically depicted in FIG. These gearboxes may be used with any engine disclosed herein, including engine 200. More detailed information regarding the gearboxes is provided below.

[0067] Engine 200 also includes a blade assembly 210 comprising a plurality of blades 212 arranged about an engine centerline axis 220. Each blade 212 has a root 248 and a tip 250, with a span defined therebetween. Blades 212 can be arranged in a variety of ways. For example, in some embodiments, they are not all equidistant from the rotating assembly.

[0068] In some embodiments, buckets 212 are mounted to a stationary frame and do not rotate relative to engine centerline axis 220, but may include mechanisms for adjusting their orientation relative to their axis 254 and / or relative to fan blades 208. For reference purposes, Figure 2 A forward direction is depicted, indicated by arrow F, which in turn defines the front and rear of the system.

[0069] like Figure 2 As depicted in FIG, in a "pull" configuration, the fan assembly 204 is located forward of the core engine 106, while the exhaust 256 is located aft of the core engine 206. Other configurations are possible and contemplated to be within the scope of the present disclosure, such as embodiments in which the engine core is located forward of the fan assembly, which may be referred to as a "push" configuration. The choice of a "pull" or "push" configuration may be consistent with the choice of mounting orientation relative to the fuselage for the intended aircraft application, and depending on whether the mounting location and orientation is a wing-mounted, fuselage-mounted, or tail-mounted configuration, some may be structurally or operationally advantageous.

[0070] A left-hand or right-hand engine configuration, which can be used in certain installations to reduce the effects of multiple engine torques on the aircraft, can be achieved by mirroring the airfoils (e.g., 208, 212) so that the fan assembly 204 rotates clockwise for one propulsion system and counterclockwise for the other propulsion system. Alternatively, an optional reversing gearbox can be provided to allow the fan blades to rotate clockwise or counterclockwise using a common gas turbine core and low-pressure turbine, i.e., to provide a left-hand or right-hand configuration as needed, so that a pair of counter-rotating engine assemblies can be provided for certain aircraft installations while eliminating the need for internal engine parts designed for counter-rotating directions.

[0071] Engine 200 also includes a gear assembly 202 including a gear set for reducing the rotational speed of fan assembly 204 relative to low-pressure turbine 244. In operation, rotating fan blades 208 are driven by low-pressure turbine 244 via gear assembly 202, causing fan blades 208 to rotate about engine centerline axis 220 and generate thrust to propel engine 200, and therefore the aircraft on which engine 200 is installed, in a forward direction F.

[0072] In some embodiments, the gear ratio of input speed to output speed is greater than 4.1. For example, in a specific embodiment, the gear ratio is in the range of 4.1-14.0, in the range of 4.5-14.0 or in the range of 6.0-14.0. In certain embodiments, the gear ratio is in the range of 4.5-12 or in the range of 6.0-11.0. Therefore, in some embodiments, the fan assembly can be configured to rotate at a speed of 400-1500rpm under cruise flight conditions, and the power turbine (e.g., low-pressure turbine) is configured to rotate at a speed of 5,000-10,000rpm under cruise flight conditions. In a specific embodiment, the fan assembly can be configured to rotate at a speed of 450-1200rpm under cruise flight conditions, and the power turbine is configured to rotate at a speed of 5,500-9,500rpm under cruise flight conditions.

[0073] It may be desirable to incorporate a pitch-changing mechanism into either or both fan blades 208 or vanes 212 so that the blades can be rotated independently or in conjunction with one another relative to a pitch axis of rotation (designated 228 and 254, respectively). Such pitch changes can be used to vary thrust and / or swirl effects under various operating conditions, including providing a thrust reversal feature that may be useful under certain operating conditions, such as when an aircraft is landing.

[0074] The size, shape, and configuration of the vanes 212 may be designed to impart counter-swirl to the fluid so that in the downstream direction behind both the fan blades 208 and the vanes 212, the fluid has a significantly reduced degree of swirl, which translates into an increased level of induced efficiency. Figure 2 As shown, blades 212 may have a shorter span than fan blades 208. For example, blades 212 may have a span that is at least 50% of the span of fan blades 208. In some embodiments, the span of the blades may be the same as or longer than the span of fan blades 208, if desired. Figure 2 As shown, blades 212 can be attached to an aircraft structure associated with engine 200, or another aircraft structure, such as a wing, pylon, or fuselage. The number of blades 212 can be less than, greater than, or equal to the number of fan blades 208. In some embodiments, the number of blades 212 is greater than two or greater than four. The size, shape, and profile of fan blades 208 can take into account the desired blade loads.

[0075] exist Figure 2 In the illustrated embodiment, annular 360-degree inlet 258 is located between fan assembly 204 and bucket assembly 210 and provides a path for incoming atmospheric air to enter core engine 206 radially inward of at least a portion of bucket assembly 210. Such a location may be advantageous for a variety of reasons, including management of icing properties and protecting inlet 258 from various objects and materials that may be encountered during operation.

[0076] exist Figure 2 In the exemplary embodiment, in addition to an open rotor or non-ducted fan assembly 204 having a plurality of fan blades 208, an optional ducted fan assembly 260 is included behind fan assembly 204, allowing engine 200 to include both ducted and non-ducted fans, both of which are used to generate thrust by moving air at ambient temperature without passing through core engine 206. Ducted fan assembly 260 is shown at approximately the same axial location as blades 212, radially inward of roots 248 of blades 212. Alternatively, ducted fan assembly 260 can be between blades 212 and core duct 262, or further forward of blades 212. Ducted fan assembly 260 can be driven by low-pressure turbine 244 or by any other suitable source of rotation and can serve as the first stage of low-pressure compressor 242 or can operate independently. Air entering inlet 258 flows through inlet duct 264 and is then split, with a portion flowing through core duct 262 and a portion flowing through fan duct 266. The fan duct 266 may incorporate one or more heat exchangers 268 and discharge to the atmosphere through a separate fixed or variable nozzle 270 located behind the bucket assembly 210, at the rear end of the fan cowl 252, and outside the engine core cowl 272. Air flowing through the fan duct 266 thus "bypasses" the engine's core and does not pass through it.

[0077] Thus, in the exemplary embodiment, engine 200 includes a non-ducted fan formed by fan blades 208, followed by a ducted fan assembly 260 that directs airflow into two concentric or non-concentric ducts 262 and 266, thereby forming a "three-flow engine architecture" with three paths for air passing through fan assembly 204. The "first flow" of engine 200 includes airflow passing outside of blade assembly 210 and / or fan cowl 252. Therefore, the first flow can be referred to as a "bypass flow" because the airflow of the first flow does not pass through core duct 262. The first flow generates the majority of the thrust of engine 200 and, therefore, can also be referred to as the "main propulsive flow." The "second flow" of engine 200 includes airflow flowing into inlet 258, through inlet duct 264, through core duct 262, and out of core nozzle 278. In this manner, the second flow can be referred to as the "core flow." The “third flow” of engine 200 includes the airflow flowing into inlet 258 , through inlet duct 264 , through fan duct 266 , and out of nozzle 270 .

[0078] As used herein, "tertiary flow" refers to a secondary air flow that is capable of increasing the energy of the flow to produce a fraction of the total thrust of an engine (e.g., engine 200). Thus, in various embodiments, a fan duct 266 having one or more heat exchangers 268 located within the flow path of the fan duct 266 may be referred to as the "tertiary flow" of a three-flow engine architecture.

[0079] The pressure ratio of the tertiary flow is higher than the pressure ratio of the main propulsive flow (i.e., the bypass flow). This thrust is generated by a dedicated nozzle (e.g., nozzle 270) or by mixing the tertiary flow with the fan flow or core flow (e.g., entering a common nozzle). In certain exemplary embodiments, the operating temperature of the airflow through the tertiary flow is less than the maximum compressor discharge temperature of the engine, and more specifically, can be less than 350 degrees Fahrenheit (e.g., less than 300 degrees Fahrenheit, such as less than 250 degrees Fahrenheit, such as less than 200 degrees Fahrenheit, and at least the same as the ambient temperature). In certain exemplary embodiments, these operating temperatures facilitate heat transfer to and from the fluid in the tertiary flow and the secondary fluid flow. Furthermore, in certain exemplary embodiments, and under takeoff conditions, or more specifically, when operating at rated takeoff power at sea level, static flight speed, and an ambient temperature of 86 degrees Fahrenheit, the airflow through the tertiary flow can contribute less than 50% of the total engine thrust (and, for example, at least 2% of the total engine thrust). Furthermore, in certain exemplary embodiments, the airflow, mixing, or exhaust characteristics of the third stream (and thereby the aforementioned exemplary percentage contribution to total thrust) can be passively adjusted during engine operation, or purposefully modified through the use of engine control features (e.g., fuel flow, motor power, variable stators, variable inlet guide vanes, valves, variable exhaust geometry, or flow characteristics) to adjust or optimize overall system performance across a wide range of potential operating conditions.

[0080] exist Figure 2 In the exemplary embodiment shown, a slidable, movable, and / or translatable plug nozzle 274 with an actuator may be included to vary the outlet area of ​​the nozzle 270. A plug nozzle is typically an annularly symmetrical device that adjusts the opening area of ​​an outlet, such as a fan flow or core flow, by axially moving the nozzle so that the clearance between the nozzle surface and a stationary structure (e.g., an adjacent wall of a duct) varies in a predetermined manner, thereby reducing or increasing the airflow space through the duct. Other suitable nozzle designs may also be employed, including those incorporating thrust reversal functionality. Such adjustable, movable nozzles may be designed to operate in conjunction with other systems (e.g., VBV, VSV, or blade pitch mechanisms) and may be designed with failure modes (e.g., fully open, fully closed, or an intermediate position) so that the nozzle 270 has a consistent "home" position to which it returns in the event of any system failure that may prevent commands from reaching the nozzle 270 and / or its actuator. In other embodiments, a static nozzle may be used.

[0081] In some embodiments, a mixing device 276 may be included in the region behind the core nozzle 278 to help mix the fan flow and the core flow, thereby improving acoustic performance by directing the core flow outward and the fan flow inward.

[0082] because Figure 2 The engine 200 shown in FIG. 1 includes an open rotor fan assembly 204 , a ducted fan assembly 260 , and a third flow so that the engine's thrust output and work distribution can be tailored to achieve specific thrust, fuel combustion, thermal management, and / or acoustic signature targets that can outperform typical ducted or non-ducted fan gas turbine propulsion assemblies of comparable thrust classes.

[0083] Operationally, engine 200 may include a control system that manages the loads of the respective open and ducted fans, and the exit area of ​​the possible variable fan nozzle, to provide different thrust, noise, cooling capacity, and other performance characteristics for various parts of the flight envelope, as well as various operating conditions associated with aircraft operation. For example, in climb mode, the ducted fans may be operated at a maximum pressure ratio, thereby maximizing the thrust capability of the flow, while in cruise mode, the ducted fans may be operated at a lower pressure ratio, improving overall efficiency by relying on thrust from the non-ducted fans. The nozzle actuator adjusts the ducted fan operating line and the overall engine fan pressure ratio, regardless of the total engine airflow. In other embodiments, static nozzles may be used to manage the loads.

[0084] As described above, the third flow (e.g., fan duct 266) may include one or more heat exchangers 268 for removing heat from various fluids used in engine operation (e.g., air-cooled oil cooler (ACOC), cooled cooling air (CCA), etc.). Compared to conventional ducted fan architectures, the heat exchanger 268 located in the third flow has reduced performance losses (e.g., fuel efficiency and thrust) by being integrated into the fan duct 266 (i.e., the third flow) because it does not affect the main thrust source, which in this case is the non-ducted fan flow. The heat exchanger can cool fluids (e.g., gearbox oil, engine sump oil), heat transfer fluids (e.g., supercritical fluids) or commercially available single-phase or two-phase fluids (supercritical CO2, EGV, Slither 800, liquid metal, etc.), engine bleed air, etc. The heat exchanger can also be composed of different sections or channels (e.g., ACOC paired with a fuel cooler) to cool different working fluids. The heat exchanger 268 may be incorporated into a thermal management system that provides heat transfer via a heat exchange fluid flowing through a network to remove heat from a source and transfer it to the heat exchanger.

[0085] Because the fan pressure ratio associated with the tertiary flow is higher for ducted fans than for non-ducted fans, the tertiary flow can provide an environment in which a more compact heat exchanger can be used than would be possible if installed outside the core shroud in a non-ducted fan flow. The fan bypass air is at a very low fan pressure ratio (FPR) (1.05 to 1.08), making it difficult to drive air through the heat exchanger. In the absence of a fan duct as described herein, a bucket or pressurized bleed air may be required to provide cooling air to and through the heat exchanger.

[0086] The third flow also offers other advantages in terms of reducing nacelle drag, enabling more aggressive nacelle closure, improved core flow particle separation, and inclement weather operations. By discharging the fan duct flow over the core cowl, this helps stimulate the boundary layer and allows for a steeper nacelle closure angle to be selected between the maximum dimensions of the engine core cowl 272 and the exhaust 256. While the closure angle is typically limited by airflow separation, boundary layer stimulation by air from the fan duct 266 discharging over the core cowl reduces airflow separation. This results in a shorter, lighter structure with less frictional surface drag.

[0087] The blade assembly can be shrouded or unshrouded (e.g. Figure 1 and 2 Although not shown, an optional annular shroud or duct may be coupled to bucket assembly 210 and positioned away from engine centerline axis 220 relative to bucket 212. In addition to the benefits of noise reduction, the duct may provide improved vibration response and structural integrity of bucket 212 by coupling bucket 212 into an assembly that forms an annular ring or one or more circumferential sectors (i.e., segments forming part of an annular ring joining two or more buckets 212). The duct may also allow for easier changes in the pitch of the bucket.

[0088] In some embodiments, the net thrust T of the engine at maximum takeoff conditions is in the range of approximately 10,000-100,000 lbf. In certain embodiments, the net thrust T of the engine at maximum takeoff conditions is in the range of approximately 12,000-35,000 lbf.

[0089] In some embodiments, an engine (e.g., engine 100 and / or engine 200) may include a counter-rotating low-pressure turbine. Figure 3-4 A schematic cross-sectional view of a counter-rotating low-pressure turbine is depicted. In particular, Figure 3 A counter-rotating turbine 300 is depicted, and Figure 4 Depicted is a counter-rotating turbine 400. The counter-rotating turbine comprises inner and outer bladed stages arranged in an alternating inner and outer configuration. In other words, the counter-rotating turbine does not include stator vanes disposed between the bladed stages.

[0090] refer to Figure 3 , counter-rotating turbine 300 includes a plurality of inner blade stages 302 and a plurality of outer blade stages 304. More specifically, counter-rotating turbine 300 includes three inner blade stages 302 coupled to and extending radially outward from an inner shaft 306 (which may also be referred to as an "inner rotor") and three outer blade stages 304 coupled to and extending radially inward from an outer shaft 308 (which may also be referred to as an "outer drum" or "outer rotor"). In this manner, counter-rotating turbine 300 may be considered a six-stage turbine.

[0091] refer to Figure 4 , counter-rotating turbine 400 includes a plurality of inner blade stages 402 and a plurality of outer blade stages 404. More specifically, counter-rotating turbine 400 includes four inner blade stages 402 coupled to and extending radially outward from inner shaft 406, and three outer blade stages 404 coupled to and extending radially inward from outer shaft 408. In this manner, counter-rotating turbine 400 can be considered a seven-stage turbine.

[0092] According to some embodiments, there are turbomachines characterized by high gear ratios. A high gear ratio gearbox refers to a gearbox having a gear ratio greater than about 4:1 to about 14:1 (or about 4.5:1 to about 12:1 in certain embodiments). For example, the engine disclosed herein may include a gearbox configured such that the output speed (i.e., the speed of the propeller) is about 400-1200 rpm under cruise flight conditions, or more specifically, 450-1000 rpm under cruise flight conditions.

[0093] Various exemplary gear assemblies are shown and described herein. In particular, Figure 5-9 Several exemplary gear assemblies that can be used with engines 100, 200 are schematically depicted. The disclosed gear assemblies can be used with any of the exemplary engines and / or any other suitable engines for which such gear assemblies may be desired. In this manner, it should be understood that the gear assemblies disclosed herein can generally operate with engines and turbomachinery, the engine having a rotating element with a plurality of rotor blades and the turbomachinery having a turbine and a shaft that can rotate with the turbine. For such an engine, a rotating element (e.g., fan assembly 104) can be driven by a shaft (e.g., a low-speed shaft) of the turbomachinery through a gear assembly.

[0094] Although the example gear assembly is shown mounted in a forward position (e.g., forward of the combustor and / or low-pressure compressor), in other embodiments, the gear assembly described herein may be mounted in a rearward position (e.g., rearward of the combustor and / or low-pressure turbine).

[0095] Various embodiments of the gear assemblies provided herein can allow for gear ratios of up to 14: 1. Still other embodiments of the gear assemblies provided herein can allow for gear ratios of at least 4.1: 1 or 4.5: 1. Still other embodiments of the gear assemblies provided herein can allow for gear ratios of 6: 1 to 12: 1.

[0096] Figure 5 There is schematically depicted a gearbox 500 that may be used, for example, with engines 100, 200. The gearbox 500 comprises a two-stage star configuration.

[0097] The first stage of gearbox 500 includes a first-stage sun gear 502, a first-stage carrier 504 housing a plurality of first-stage planetary gears, and a first-stage ring gear 506. First-stage sun gear 502 can be coupled to a low-speed shaft 508, which in turn is coupled to the engine's low-pressure turbine. First-stage sun gear 502 can mesh with the first-stage planetary gears, which in turn mesh with the first-stage ring gear. First-stage carrier 504 can be fixed against rotation by a support member 510.

[0098] The second stage of gearbox 500 includes a second-stage sun gear 512, a second-stage carrier 514 that houses a plurality of second-stage planetary gears, and a second-stage ring gear 516. Second-stage sun gear 512 can be coupled to shaft 518, which in turn is coupled to first-stage ring gear 506. Second-stage carrier 514 can be fixed against rotation by support members 520. Second-stage ring gear 516 can be coupled to fan shaft 522.

[0099] In some embodiments, each stage of gearbox 500 may include five spider gears. In other embodiments, gearbox 500 may include fewer or more than five spider gears in each stage. In some embodiments, the first-stage carrier may include a different number of spider gears than the second-stage carrier. For example, the first-stage carrier may include five spider gears and the second-stage carrier may include three spider gears, or vice versa.

[0100] In some embodiments, the radius R of the gearbox 500 l It can be about 16-19 inches. In other embodiments, the radius R of the gearbox 500 l It can be about 22-24 inches. In other embodiments, the radius R of the gearbox 500 l Can be smaller than 16 inches or larger than 24 inches.

[0101] Figure 6A gearbox 600 that can be used with the engine disclosed herein is schematically depicted. Gearbox 600 includes a single-stage planetary configuration. Gearbox 600 includes a sun gear 602, a carrier 604 that accommodates a plurality of planetary gears (e.g., 3-5 planetary gears), and a ring gear 606. Sun gear 602 can mesh with the planetary gears, and the planetary gears can mesh with ring gear 606. Sun gear 602 can be coupled to a low-speed shaft 608, which in turn is coupled to the engine's low-pressure turbine. Carrier 604 can be fixed against rotation by a support member 610. Ring gear 606 can be coupled to a fan shaft 612.

[0102] In some embodiments, the radius R2 of the gearbox 600 can be approximately 18-23 inches. In other embodiments, the radius R2 of the gearbox 600 can be less than 18 inches or greater than 23 inches.

[0103] Figure 7 Another exemplary gearbox 700 is schematically depicted. Gearbox 700 includes a single-stage planetary configuration. Gearbox 700 includes a sun gear 702, a carrier 704 that accommodates a plurality of planetary gears (e.g., 3-5 planetary gears), and a ring gear 706. Sun gear 702 can mesh with the planetary gears, and the planetary gears can mesh with ring gear 706. Sun gear 702 can be coupled to a low-speed shaft 708, which in turn is coupled to the low-pressure turbine of the engine. Carrier 704 can be fixed against rotation by a support member 710. Ring gear 706 can be coupled to a fan shaft 712.

[0104] The gearbox 700 can include a radius R3. In some embodiments, the radius R3 of the gearbox 700 can be approximately 10-13 inches. In other embodiments, the radius R3 of the gearbox 700 can be less than 10 inches or greater than 13 inches.

[0105] Figure 8 A gearbox 800 that may be used with the engines disclosed herein is schematically depicted. The gearbox 800 includes a two-stage configuration, where the first stage is a star configuration and the second stage is a planetary configuration.

[0106] The first stage of gearbox 800 includes a first-stage sun gear 802, a first-stage spider carrier 804 containing a plurality of first-stage spider gears (e.g., 3-5 spider gears), and a first-stage ring gear 806. First-stage sun gear 802 can mesh with the first-stage spider gears, and the first-stage spider gears can mesh with the first-stage ring gear 806. First-stage sun gear 802 can be coupled to a higher-speed shaft 808 of the low-spool, which in turn is coupled to the inner blades of the engine's low-pressure turbine. First-stage spider carrier 804 can be fixed against rotation by a support member 810.

[0107] The second stage of gearbox 800 includes a second-stage sun gear 812, a second-stage planet carrier 814 containing a plurality of second-stage planet gears (e.g., 3-5 planet gears), and a second-stage ring gear 816. Second-stage sun gear 812 can mesh with the second-stage planet gears. Second-stage planet carrier 814 can be coupled to first-stage ring gear 806. Second-stage sun gear 812 can be coupled to a lower speed shaft 818 of the low-spool, which in turn is coupled to the outer blades of the engine's low-pressure turbine. Second-stage planet carrier 814 can be coupled to first-stage ring gear 806. Second-stage planet carrier 814 can also be coupled to fan shaft 820. Second-stage ring gear 816 can be fixed against rotation by a support member 822.

[0108] In some embodiments, each stage of gearbox 800 may include three planetary gears. In other embodiments, gearbox 800 may include fewer or more than three planetary gears per stage. In some embodiments, the first-stage carrier may include a different number of planetary gears than the second-stage carrier. For example, the first-stage carrier may include five planetary gears, and the second-stage carrier may include three planetary gears, or vice versa.

[0109] Because the first stage of gearbox 800 is coupled to the higher-speed shaft 808 of the low spool, and the second stage of gearbox 800 is coupled to the lower-speed shaft 818 of the low spool, the gear ratio of the first stage of gearbox 800 can be greater than the gear ratio of the second stage of gearbox. For example, in certain embodiments, the first stage of gearbox 800 can include a gear ratio of 4.1-14, and the second stage of gearbox 800 can include a gear ratio that is less than the gear ratio of the first stage of gearbox. In a particular embodiment, the first stage of gearbox 800 can include a gear ratio of 7, and the second stage of gearbox 800 can include a gear ratio of 6.

[0110] In some embodiments, the engine including gearbox 800 can be configured such that the higher-speed shaft 808 provides approximately 50% of the power to gearbox 800 and the lower-speed shaft 818 provides approximately 50% of the power to gearbox 800. In other embodiments, the engine including gearbox 800 can be configured such that the higher-speed shaft 808 provides approximately 60% of the power to gearbox 800 and the lower-speed shaft 818 provides approximately 40% of the power to gearbox 800.

[0111] In some embodiments, the radius R4 of the gearbox 800 can be approximately 18-22 inches. In other embodiments, the radius R4 of the gearbox 800 can be less than 18 inches or greater than 22 inches.

[0112] Figure 9Gearbox 900 is depicted, which can be used, for example, with the engines disclosed herein (e.g., engines 100, 200). Gearbox 900 is configured as a compound planetary gearbox. Gearbox 900 includes a sun gear 902 and a planetary carrier 904, which includes a plurality of compound planetary gears having one or more first portions 906 and one or more second portions 908. Gearbox 900 also includes a ring gear 910. Sun gear 902 can also mesh with the first portion 906 of the planetary gears. The planetary carrier can be fixed against rotation via a support member 914. The second portion 908 of the planetary gears can mesh with the ring gear 910. Sun gear 902 can be coupled to a low-pressure turbine via a turbine shaft 912. Ring gear 910 can be coupled to a fan shaft 916.

[0113] Each embodiment of the turbomachinery disclosed herein utilizes an open rotor (also known as a "non-ducted propeller") and a gearbox. Open rotor engines that utilize a geared transmission present unique challenges. One of these challenges is finding the right balance between engine performance, engine efficiency, propeller size, and / or acoustics (e.g., noise). For example, if the propeller is larger than a certain diameter, the engine performance may be satisfactory, but the engine will not fit on the airframe. If the fan is smaller, the engine performance may not be ideal. If other criteria are met, the engine may still be too noisy to comply with regulations and / or provide a poor experience for those flying in the aircraft.

[0114] Starting from this basis, the inventors set out to calculate the various requirements for non-ducted propulsors. In the process of developing the above-mentioned embodiments of turbomachinery combining non-ducted propulsors and gearboxes, the inventors unexpectedly discovered that a good approximation of the design of an integrally geared non-ducted propulsor engine can be made using only relatively few engine parameters. Among other things, this development was based on an understanding of the relationship between the propeller, LPT and gearbox. From this initial understanding and other developments that were a byproduct of studying various engine configurations (including the configuration disclosed herein), the inventors ultimately discovered that a good approximation could be made to overall engine performance, efficiency, size and acoustics based on the relationship between the fan diameter, the number of LPT stages and the gear ratio of the gearbox. The inventors refer to this newly discovered relationship as Engine Correlated Parameters (ECP).

[0115] ECP can be highly beneficial. For example, it can provide an engine with adequate performance, efficiency, size, and acoustics. Consequently, ECP can improve the process of developing non-ducted engines, which can ultimately lead to improved turbomachinery. For example, by using ECP to define engine architectures, a better understanding of the affected systems or subsystems, reliability or feasibility, and / or the range of possible architectures can be achieved because ECP imposes practical constraints on the considered design.

[0116] The ECP is defined by the following relationship: D / N / GR, where D is the blade tip diameter measured in feet, where N is the number of low-pressure turbine stages, and where GR is the gear ratio of the gearbox. For example, the ECP can be used to construct an engine with higher efficiency and lower noise than a typical engine. In some cases, the engines disclosed herein that include the ECP are lighter, have higher propulsion efficiency, have higher engine efficiency, and / or are quieter than a typical engine. The ECP thus provides an improved turbomachinery engine and / or can help simplify one or more complexities of developing a geared open rotor engine.

[0117] Specifically, the inventors of the present disclosure have recognized that for a single non-ducted rotor engine (e.g., engines 100, 200), an ECP between about 0.17 feet and about 0.83 feet, which is equal to the propeller blade tip diameter (in feet) divided by the gear ratio divided by the number of low-pressure turbine stages, provides improved results during operation. More specifically, in at least some examples, a single non-ducted rotor engine with an ECP between about 0.17 feet and about 0.63 feet can provide desirable results during operation. For simplicity, ECP values ​​can be expressed as unitless values ​​(i.e., omitting the "feet"). For example, an ECP of 0.19 feet can be simply expressed as 0.19.

[0118] For example, in one example, a geared single non-ducted rotor engine may include an 18-foot propeller blade diameter, a 10:1 gear ratio, and an LPT stage number equal to 3. In such an example, the ECP is equal to 0.6.

[0119] As another example, a geared single non-ducted rotor engine may include a propeller blade diameter of 14 feet, may have a gear ratio of 7:1, and a number of LPT stages equal to 5. In this example, the ECP is equal to 0.4.

[0120] Figure 10-16 A table with additional exemplary engines and their associated ECPs is depicted. Figure 10-16 The engines disclosed in the table (and elsewhere in this disclosure) are examples only and are not intended to limit the scope of this disclosure.

[0121] Figure 10 Various engines (i.e., engines 1-10) are depicted, each including a fan diameter of 10 feet. The LPT stages of each engine 1-10 are in the range of 3 to 8. The gear ratios of each engine 1-10 are in the range of 4 to 12. The ECP of each engine 1-10 is in the range of 0.17 to 0.83.

[0122] Figure 11Various engines (i.e., engines 11-20) are depicted, each including a fan diameter of 11 feet. The LPT stages of each engine 11-20 are in the range of 3-8. The gear ratios of each engine 11-20 are in the range of 4-12. The ECP of each engine 11-20 is in the range of 0.17 to 0.83.

[0123] Figure 12 Various engines (i.e., engines 21-30) are depicted, including a fan diameter of 12 feet. The LPT stages of engines 21-30 are in the range of 3 to 7. The gear ratios of engines 21-30 are in the range of 4 to 11. The ECPs of engines 21-30 are in the range of 0.17 to 0.83.

[0124] Figure 13 Various engines (i.e., engines 31-40) are depicted, each including a fan diameter of 13 feet. The LPT stages of each engine 31-40 are in the range of 3-8. The gear ratios of each engine 31-40 are in the range of 4-11. The ECP of each engine 31-40 is in the range of 0.17 to 0.83.

[0125] Figure 14 Various engines (i.e., engines 41-50) are depicted, including a fan diameter of 14 feet. The LPT stages of engines 41-50 are in the range of 3 to 7. The gear ratios of engines 41-50 are in the range of 5 to 12. The ECPs of engines 41-50 are in the range of 0.17 to 0.83.

[0126] Figure 15 Various engines (i.e., engines 51-60) are depicted, each including a fan diameter of 15 feet. The LPT stages of each engine 51-60 are in the range of 3-8. The gear ratios of each engine 51-60 are in the range of 4-12. The ECP of each engine 51-60 is in the range of 0.17 to 0.83.

[0127] Figure 16 Various engines (i.e., engines 61-70) are depicted, including a fan diameter of 16 feet. The LPT stages of engines 61-70 are in the range of 3 to 8. The gear ratios of engines 61-70 are in the range of 4 to 12. The ECP of engines 61-70 is in the range of 0.17 to 0.83.

[0128] It should be noted that although the exemplary engine 1-70 includes fan diameters that are integers (e.g., 10 ft, 11 ft, 12 ft, 14 ft, etc.), the engine may include fan diameters that are not integers. For example, the fan may include a diameter of 11.5 ft, 12.75 ft, 13.2 ft, 14.1 ft, 15.56 ft, or 16.3 ft.

[0129] The inventors have also discovered a relationship between ECP and propeller disc loading. Disc loading is the power measured in horsepower (HP) divided by the square feet (ft) of propeller under takeoff flight conditions. 2 ) is the swept area of ​​the propeller measured in units of . For example, the inventors have found that an engine with an ECP of 0.17-0.83 and a power of 60-180 HP / ft 2 The propeller disc load between is advantageous, as Figure 17 More specifically, the inventors have found that the engine has an ECP of 0.17-0.63 and a power output of 60-180 HP / ft at takeoff flight conditions. 2 The propeller disc load between is particularly advantageous, as Figure 18 For example, compared to a typical engine with an ECP in the range of 0.17-0.83 ft (or 0.17-0.63 ft) and 60-180 HP / ft at takeoff flight conditions 2 A propeller disc loaded engine may advantageously provide improved engine performance and efficiency while also meeting acoustic requirements.

[0130] In some examples, the propulsion assembly of the engine has a power output of 60-180 HP / ft at takeoff flight conditions. 2 In some cases, under takeoff flight conditions, the net efficiency of the propeller assembly is in the range of approximately 0.62-0.75 and the propeller assembly has a disc loading of 60-100 HP / ft. 2 In a specific example, under takeoff flight conditions, the net efficiency of the propeller assembly is in the range of approximately 0.57-0.67, and the disc loading of the propeller assembly is in the range of 160-180 HP / ft 2 In a specific example, under takeoff flight conditions, the net efficiency of the propeller assembly is in the range of approximately 0.58-0.72, and the disc loading of the propeller assembly is in the range of 100-160 HP / ft 2 The net efficiency varies with disk load, so that for 60 to 180 HP / ft 2For disk loads of 0.65 to 0.57, the lower limit can be assumed to vary linearly, and for 60 to 180 HP / ft 2 For disc loads of 0.75 to 0.67, it can be assumed that the upper limit varies linearly. Referring to Table 1 below, the relationship between net efficiency and disc load is shown:

[0131]

[0132] As previously mentioned, turbomachinery engines (e.g., engines 100, 200) include many variables and factors that affect their performance and / or operation. The interactions between the various components can make developing or selecting a component particularly difficult, especially when each component is at a different stage of completion. For example, one or more components may be nearing completion, while one or more other components may be in an initial or preliminary stage where only one (or several) parameters are known. In addition, each component often undergoes more than one change during development, which typically lasts for many years (e.g., 5-15 years). These complex and intricate individual and collective development processes can be both cumbersome and inefficient. For at least these reasons, there is a need for devices and methods that can provide good estimates, not only of the basic configuration or dimensions required to achieve the desired performance benefits, but also of the penalties or adaptations that reflect other areas in order to achieve the desired benefits.

[0133] According to another aspect of the present disclosure, the ECP can also provide a particularly useful indication of the efficiency and effectiveness of an engine during initial development, for example, as a tool for accepting or rejecting a particular configuration. Thus, the ECP can be used, for example, to guide engine development. For example, the ECP can be used to quickly and accurately determine the suitability of a particular engine without requiring an individual or team to complete the tedious and time-consuming process of fully developing an engine. Thus, the ECP can improve the process of developing turbomachinery engines.

[0134] This written description uses examples to disclose the technology, including the best mode, and also to enable anyone 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. If these other examples include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements that do not differ substantially from the literal language of the claims, then these other examples are intended to fall within the scope of the claims.

[0135] Further aspects of the present disclosure are provided by the subject matter of the following examples:

[0136] Example 1. A turbomachinery engine comprising a non-ducted rotor assembly, a low-pressure turbine, a low-pressure shaft, a gearbox, and engine-related parameters. The non-ducted rotor assembly includes a plurality of rotor blades arranged in a single row and defining a blade diameter. The low-pressure turbine includes a number of stages defined by the plurality of rows of rotating blades of the low-pressure turbine. The low-pressure shaft is coupled to the low-pressure turbine. The gearbox includes an input, an output, and a gear ratio. The input of the gearbox is coupled to the low-pressure shaft, and the output of the gearbox is coupled to the non-ducted rotor assembly. The engine-related parameters are greater than 0.17 and less than 0.83. The engine-related parameters are equal to D / N / GR, where D is the blade tip diameter measured in feet, N is the number of stages of the low-pressure turbine, and GR is the gear ratio of the gearbox.

[0137] Example 2. The turbomachinery engine of any example herein, wherein the engine-related parameter is greater than 0.17 and less than 0.63.

[0138] Example 3. The turbomachinery engine of any example herein, wherein the gear ratio is in the range of 4:1 to 12:1.

[0139] Example 4. The turbomachinery engine according to any example herein, wherein the number of low-pressure turbine stages is between 3 and 8.

[0140] Example 5. The turbomachinery engine according to any example herein, wherein the number of low-pressure turbine stages is between 4 and 6.

[0141] Example 6. The turbomachinery engine of any example herein, wherein the non-ducted rotor assembly includes 8-20 rotor blades.

[0142] Example 7. The turbomachinery engine of any example herein, wherein the non-ducted rotor assembly includes 8-14 rotor blades.

[0143] Example 8. The turbomachinery engine of any example herein, further comprising a plurality of non-rotating exit guide vanes disposed behind the rotor blades of the non-ducted rotor assembly.

[0144] Example 9. The turbomachinery engine of any example herein, wherein the blade tip diameter is in the range of 10-16 feet.

[0145] Example 10. The turbomachinery engine of any example herein, wherein the blade tip diameter is in the range of 11-14 feet.

[0146] Example 11. The turbomachinery engine of any example herein, wherein the blade tip diameter is in the range of 12-13 feet.

[0147] Example 12. The turbomachinery engine of any example herein, wherein the non-ducted rotor assembly is configured to rotate at a speed less than 1000 rpm and greater than 450 rpm under cruise flight conditions.

[0148] Example 13. The turbomachinery engine of any example herein, wherein the turbomachinery engine is configured to operate at a cruise flight condition greater than Mach 0.6 and less than Mach 0.9.

[0149] Example 14. The turbomachinery engine of any example herein, further comprising a net thrust in the range of 10,000-100,000 lbf at maximum takeoff conditions.

[0150] Example 15. The turbomachinery engine of any example herein, further comprising a net thrust in the range of 20,000-35,000 lbf at maximum takeoff conditions.

[0151] Example 16. The turbomachinery engine of any example herein, wherein the non-ducted rotor assembly comprises 60-180 HP / ft at takeoff flight conditions. 2 range of disc loads.

[0152] Example 17. The turbomachinery engine of any example herein, wherein the fan comprises a firmness in the range of 0.5-1.0.

[0153] Example 18. The turbomachinery engine of any example herein, wherein the fan comprises a firmness in the range of 0.6-1.0.

[0154] Example 19. The turbomachinery engine of any example herein, wherein the fan comprises a firmness in the range of 1.1-1.5.

[0155] Example 20. The turbomachinery engine of any example herein, wherein the fan comprises a firmness in the range of 1.1-1.3.

[0156] Example 21. A single non-ducted rotor turbomachinery engine includes a low-pressure turbine, a low-pressure spool, a gearbox, and a non-ducted rotor assembly. The low-pressure turbine defines a number of stages. The low-pressure spool is coupled to the low-pressure turbine. The gearbox defines a gear ratio. The non-ducted rotor assembly includes a plurality of rotor blades defining a blade tip diameter, and the non-ducted rotor assembly is coupled to the low-pressure spool via the gearbox. The single non-ducted rotor turbomachinery engine defines an engine-related parameter equal to the blade tip diameter measured in feet divided by the number of low-pressure turbine stages divided by the gear ratio of the gearbox, and the engine-related parameter is greater than 0.17 and less than 0.83.

[0157] Example 22. A single non-ducted rotor turbomachinery engine according to any example herein, wherein the engine-related parameter is greater than 0.17 and less than 0.63.

[0158] Example 23. A single non-ducted rotor turbomachinery engine according to any example herein, wherein the gear ratio is between 4:1 and 12:1.

[0159] Example 24. A single non-ducted rotor turbomachinery engine according to any example herein, wherein the number of stages is between 3 and 8.

[0160] Example 25. A single non-ducted rotor turbomachinery engine according to any example herein, wherein the number of low-pressure turbine stages is between 4 and 6.

[0161] Example 26. The single non-ducted rotor turbomachinery engine of any example herein, wherein the non-ducted rotor assembly comprises 8-20 rotor blades.

[0162] Example 27. The single non-ducted rotor turbomachinery engine of any example herein, wherein the non-ducted rotor assembly comprises 8-14 rotor blades.

[0163] Example 28. The single non-ducted rotor turbomachinery engine of any example herein, further comprising a plurality of non-rotating exit guide vanes disposed behind the rotor blades of the non-ducted rotor assembly.

[0164] Example 29. A single non-ducted rotor turbomachinery engine according to any example herein, wherein the blade tip diameter is in the range of 10-16 feet.

[0165] Example 30. The single non-ducted rotor turbomachinery engine of any example herein, wherein the blade tip diameter is in the range of 11-14 feet.

[0166] Example 31. A single non-ducted rotor turbomachinery engine according to any example herein, wherein the blade tip diameter is in the range of 12-13 feet.

[0167] Example 32. A single non-ducted rotor turbomachinery engine according to any example herein, wherein the non-ducted rotor assembly is configured to rotate at a speed less than 1000 rpm and greater than 450 rpm under cruise flight conditions.

[0168] Example 33. A single non-ducted rotor turbomachinery engine according to any example herein, wherein the single non-ducted rotor turbomachinery engine is configured to operate at a cruise flight condition greater than Mach 0.7 and less than Mach 0.9.

[0169] Example 34. The single non-ducted rotor turbomachinery engine of any example herein, further comprising a net thrust in the range of 10,000-100,000 lbf at maximum takeoff conditions.

[0170] Example 35. The single non-ducted rotor turbomachinery engine of any example herein, further comprising a net thrust in the range of 20,000-35,000 lbf at maximum takeoff conditions.

[0171] Example 36. A single non-ducted rotor turbomachinery engine according to any example herein, wherein the non-ducted rotor assembly comprises 60-180 HP / ft at takeoff flight conditions. 2 range of disc loads.

[0172] Example 37. A single non-ducted rotor turbomachinery engine according to any example herein, wherein the fan comprises a firmness in the range of 0.5-1.0.

[0173] Example 38. A single non-ducted rotor turbomachinery engine according to any example herein, wherein the fan comprises a firmness in the range of 0.6-1.0.

[0174] Example 39. A single non-ducted rotor turbomachinery engine according to any example herein, wherein the fan comprises a firmness in the range of 1.1-1.5.

[0175] Example 40. A single non-ducted rotor turbomachinery engine according to any example herein, wherein the fan comprises a firmness in the range of 1.1-1.3.

[0176] Example 41. A turbomachinery engine comprising a non-ducted propeller assembly, a low-pressure compressor, a low-pressure turbine, a low-pressure shaft, a high-pressure compressor, a high-pressure turbine, a high-pressure shaft, a gearbox, and engine-related parameters. The non-ducted propeller assembly includes a plurality of propeller blades arranged in a single row and defining a blade diameter, and the blade tip diameter is in the range of 8-16 feet. The low-pressure compressor includes one or more LPC rotors defining a number of LPC stages. The low-pressure turbine includes a plurality of LPT rotors defining a number of LPT stages, and the number of LPT stages is in the range of 3-8. The low-pressure shaft is coupled to the low-pressure compressor and the low-pressure turbine. The high-pressure compressor includes a plurality of HPC rotors defining a number of HPC stages. The high-pressure turbine includes one or more HPT rotors defining a number of HPT stages. The high-pressure shaft is coupled to the high-pressure compressor and the high-pressure turbine. The gearbox includes an input, an output, and a gear ratio. The input is coupled to the low-pressure shaft and is configured to rotate at a first speed. The output is coupled to the non-ducted propeller assembly and is configured to rotate at a second speed less than the first speed. The gear ratio is defined by the ratio of the first rotational speed to the second rotational speed and is in the range of 4 to 12. The engine-related parameter is in the range of 0.17 to 0.83, where the engine-related parameter is equal to D / N / GR, where D is the blade tip diameter of the propeller blade measured in feet, N is the number of LPT stages of the low pressure turbine, and GR is the gear ratio of the gearbox.

[0177] Example 42. The turbomachinery engine of any example herein, wherein the engine-related parameter is within the range of 0.17-0.63.

[0178] Example 43. The turbomachinery engine of any example herein, wherein the non-ducted propeller assembly comprises 8-20 propeller blades.

[0179] Example 44. The turbomachinery engine of any example herein, wherein the non-ducted propeller assembly comprises 8-14 propeller blades.

[0180] Example 45. The turbomachinery engine of any example herein, further comprising a plurality of stationary exit guide vanes disposed behind the propeller blades.

[0181] Example 46. The turbomachinery engine of any example herein, wherein the blade tip diameter is in the range of 8-14 feet, 11-14 feet, or 10-16 feet.

[0182] Example 47. The turbomachinery engine of any example herein, wherein the blade tip diameter is in the range of 12-13 feet.

[0183] Example 48. The turbomachinery engine of any example herein, wherein the non-ducted propeller assembly is configured to rotate within a range of 450-1000 rpm under cruise flight conditions.

[0184] Example 49. The turbomachinery engine of any example herein, wherein the turbomachinery engine is configured to operate in the Mach 0.6-0.9 range under cruise flight conditions.

[0185] Example 50. The turbomachinery engine of any example herein, further comprising a net thrust in the range of 10,000-100,000 lbf at maximum takeoff conditions.

[0186] Example 51. The turbomachinery engine of any example herein, further comprising a net thrust in the range of 20,000-35,000 lbf at maximum takeoff conditions.

[0187] Example 52. A turbomachinery engine according to any example herein, wherein the propeller assembly comprises 60-180 HP / ft at takeoff flight conditions. 2 range of disc loads.

[0188] Example 53. The turbomachinery engine of any example herein, wherein the gearbox is an epicyclic gearbox comprising a sun gear, a plurality of planet gears, and a ring gear, wherein the sun gear is an input and wherein the ring gear is an output.

[0189] Example 54. A turbomachinery engine according to any example herein, wherein the gearbox is an epicyclic gearbox comprising a sun gear, a plurality of planet gears, and a ring gear, wherein the sun gear is an input, wherein the planet gears are coupled to a planet carrier, and wherein the planet carrier is an output.

[0190] Example 55. The turbomachinery engine of any example herein, wherein the gearbox is a multi-stage gearbox.

[0191] Example 56. The turbomachinery engine of any example herein, wherein the gearbox is a two-stage gearbox.

[0192] Example 57. The turbomachinery engine of any example herein, wherein the gearbox is a compound gearbox.

[0193] Example 58. A turbomachinery engine according to any example herein, wherein the number of LPC stages is in the range of 1-3.

[0194] Example 59. A turbomachinery engine according to any example herein, wherein the number of HPC stages is in the range of 10-11.

[0195] Example 60. A turbomachinery engine according to any example herein, wherein the number of HPT stages is two.

[0196] Example 61. A turbomachinery engine according to any example herein, wherein the number of LPT stages is four.

[0197] Example 62. A turbomachinery engine according to any example herein, wherein the low pressure turbine is a counter-rotating low pressure turbine, wherein the LPT rotor comprises a plurality of inner rotors and a plurality of outer rotors arranged in an alternating configuration.

[0198] Example 63. A turbomachinery engine according to any example herein, wherein the fan comprises a firmness in the range of 0.5-1.0.

[0199] Example 64. A turbomachinery engine according to any example herein, wherein the fan comprises a firmness in the range of 0.6-1.0.

[0200] Example 65. The turbomachinery engine of any example herein, wherein the fan comprises a firmness in the range of 1.1-1.5.

[0201] Example 66. The turbomachinery engine of any example herein, wherein the fan comprises a firmness in the range of 1.1-1.3.

[0202] Example 67. The turbomachinery engine of any example herein, wherein the turbomachinery engine comprises a third flow.

[0203] Example 68. A turbomachinery engine according to any example herein, wherein the propeller assembly has a thrust capacity of 60-180 HP / ft at takeoff flight conditions. 2 of disc loading and a net efficiency in the range of 0.57-0.75.

[0204] Example 69. The turbomachinery engine of any example herein, wherein the net efficiency of the propeller assembly is in the range of 0.64-0.75 and the disc loading of the propeller assembly is in the range of 60-80 HP / ft at takeoff flight conditions. 2 within the range.

[0205] Example 70. The turbomachinery engine of any example herein, wherein at takeoff flight conditions, the net efficiency of the propeller assembly is in the range of 0.57-0.68, and the disc loading of the propeller assembly is in the range of 160-180 HP / ft 2 within the range.

[0206] Example 71. A turbomachinery engine according to any example herein, wherein at takeoff flight conditions, the net efficiency of the propeller assembly is in the range of 0.58-0.72 and the disc loading of the propeller assembly is in the range of 100-160 HP / ft 2 within the range.

Claims

1. A turbomachinery engine, characterized in that: include: a non-ducted rotor assembly comprising a plurality of rotor blades arranged in a single row and defining a blade tip diameter; a low-pressure turbine comprising a number of stages defined by a plurality of rows of rotating blades of the low-pressure turbine; a low-pressure shaft coupled to the low-pressure turbine; a gearbox comprising an input, an output, and a gear ratio, wherein the input of the gearbox is coupled to the low-pressure shaft, and wherein the output of the gearbox is coupled to the non-ducted rotor assembly; and an engine-related parameter greater than 0.17 and less than 0.83, wherein the engine-related parameter is equal to D / N / GR, wherein D is the blade tip diameter measured in feet, wherein N is the number of stages of the low-pressure turbine, and wherein GR is the gear ratio of the gearbox.

2. The turbomachinery engine according to claim 1, wherein: The engine-related parameters are greater than 0.17 and less than 0.

63.

3. The turbomachinery engine according to claim 1, wherein: The gear ratio is in the range of 4:1 to 12:

1.

4. The turbomachinery engine according to claim 1, wherein: The number of stages of the low-pressure turbine is between 3 and 8.

5. The turbomachinery engine according to claim 1, wherein: The number of stages of the low-pressure turbine is between 4 and 6.

6. The turbomachinery engine according to claim 1, wherein: The non-ducted rotor assembly includes 8-20 rotor blades.

7. The turbomachinery engine according to claim 1, wherein: The non-ducted rotor assembly includes 8-14 rotor blades.

8. The turbomachinery engine according to claim 1, wherein: wherein the non-ducted rotor assembly comprises 60-180 HP / ft at takeoff flight conditions 2 range of disc loads.

9. The turbomachinery engine according to claim 1, wherein: Further comprising a third stream.

10. A single non-ducted rotor turbomachinery engine, characterized in that: include: a low-pressure turbine, wherein the low-pressure turbine has a limited number of stages; a low-pressure spool coupled to the low-pressure turbine; a gearbox defining a gear ratio; as well as a non-ducted rotor assembly including a plurality of rotor blades defining a blade tip diameter, the non-ducted rotor assembly coupled to the low pressure spool via the gearbox, wherein the single non-ducted rotor turbomachinery engine defines an engine-related parameter equal to the blade tip diameter measured in feet divided by the number of stages of the low-pressure turbine divided by the gear ratio of the gearbox, and wherein the engine-related parameter is greater than 0.17 and less than 0.

83.

11. The single non-ducted rotor turbomachinery engine according to claim 10, wherein: The engine-related parameters are greater than 0.17 and less than 0.

63.

12. The single non-ducted rotor turbomachinery engine according to claim 10, wherein: The gear ratio is between 4:1 and 12:

1.

13. The single non-ducted rotor turbomachinery engine according to claim 10, wherein: The number of stages is between 3 and 8.

14. The single non-ducted rotor turbomachinery engine according to claim 10, wherein: The number of stages of the low-pressure turbine is between 4 and 6.

15. The single non-ducted rotor turbomachinery engine according to claim 10, wherein: The non-ducted rotor assembly includes 8-20 rotor blades.

16. The single non-ducted rotor turbomachinery engine according to claim 10, wherein: The non-ducted rotor assembly includes 8-14 rotor blades.

17. The single non-ducted rotor turbomachinery engine according to claim 10, wherein: wherein the non-ducted rotor assembly comprises 60-180 HP / ft at takeoff flight conditions 2 range of disc loads.

18. The single non-ducted rotor turbomachinery engine of claim 10, wherein: Further comprising a third stream.

19. A turbomachinery engine, characterized in that: include: a non-ducted propeller assembly comprising a plurality of propeller blades arranged in a single row and defining a blade tip diameter, wherein the blade tip diameter is in the range of 8-16 feet; a low-pressure compressor comprising one or more LPC rotors defining a number of LPC stages; a low-pressure turbine comprising a plurality of LPT rotors defining a number of LPT stages, wherein the number of LPT stages is in the range of 3-8; a low-pressure shaft coupled to the low-pressure compressor and the low-pressure turbine; a high pressure compressor comprising a plurality of HPC rotors defining a number of HPC stages; a high-pressure turbine comprising one or more HPT rotors defining a number of HPT stages; a high-pressure shaft coupled to the high-pressure compressor and the high-pressure turbine; a gearbox comprising an input, an output, and a gear ratio, wherein the input is coupled to the low-pressure shaft and configured to rotate at a first rotational speed, wherein the output is coupled to the non-ducted propeller assembly and configured to rotate at a second rotational speed less than the first rotational speed, and wherein the gear ratio is defined by a ratio of the first rotational speed to the second rotational speed and is in the range of 4-12; as well as an engine-related parameter in the range of 0.17-0.83, wherein the engine-related parameter is equal to D / N / GR, wherein D is the blade tip diameter of the propeller blade measured in feet, wherein N is the number of LPT stages of the low pressure turbine, and wherein GR is the gear ratio of the gearbox.

20. The turbomachinery engine according to claim 19, wherein: The engine-related parameters are in the range of 0.17-0.

63.

21. The turbomachinery engine according to claim 19, wherein: The non-ducted propeller assembly includes 8-20 propeller blades.

22. The turbomachinery engine according to claim 19, wherein: The non-ducted propeller assembly includes 8-14 propeller blades.

23. The turbomachinery engine according to claim 19, wherein: Further included is a plurality of stationary exit guide vanes disposed behind the propeller blades.

24. The turbomachinery engine according to claim 19, wherein: wherein the blade tip diameter is in the range of 8-14 feet, 11-14 feet, or 10-16 feet.

25. The turbomachinery engine according to claim 19, wherein: Further comprising a third stream.

26. The turbomachinery engine according to claim 19, wherein: wherein the non-ducted propulsion assembly comprises 60-180HP / ft under takeoff flight conditions 2 range of disc loads.

27. The turbomachinery engine according to claim 26, wherein: The non-ducted thruster assembly has a net efficiency in the range of 0.57-0.75 under takeoff flight conditions.

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