Fan blade assembly with mid-span shroud

Abstract

An engine includes a fan section comprising multiple fan blade assemblies. Each fan blade assembly includes a mid-span shroud separating an inner portion and an outer portion. The outer pitch of the outer portion is variable relative to the inner pitch of the inner portion.

Description

Technical Field

[0001] This disclosure relates to fan blade assemblies, such as fan blade assemblies for variable pitch fans used in engines. Background Technology

[0002] Gas turbine engines typically consist of a turbine and a rotor assembly. Gas turbine engines (such as turbofan engines) are used for aircraft propulsion. Some turbofan engines include a fan configured to generate forward thrust during flight operations. During other operations, the fan can be configured to generate reverse thrust to, for example, decelerate an aircraft incorporating a turbofan engine, during landing operations. Improving the fan of a turbofan engine to facilitate reverse thrust would be welcome in the art. Attached Figure Description

[0003] The complete and feasible disclosure of this disclosure, including its best mode, is set forth in the specification with reference to the accompanying drawings, for those skilled in the art, wherein:

[0004] Figure 1 This is a cross-sectional view of a gas turbine engine according to an exemplary aspect of this disclosure.

[0005] Figure 2 Based on exemplary aspects of this disclosure Figure 1 A schematic front view of the fan section of a gas turbine engine.

[0006] Figure 3 Based on exemplary aspects of this disclosure Figure 2 An end view of the fan blade assembly in the fan section.

[0007] Figure 4 Based on exemplary aspects of this disclosure Figure 2 A three-dimensional view of the fan blade assembly in the fan section.

[0008] Figure 5 Based on exemplary aspects of this disclosure Figure 4 End view of the fan blade assembly.

[0009] Figure 6 Based on exemplary aspects of this disclosure Figure 2 A three-dimensional view of the fan blade assembly in the fan section.

[0010] Figure 7 Based on exemplary aspects of this disclosure Figure 6 End view of the fan blade assembly.

[0011] Figure 8 This is a schematic diagram of a fan blade assembly according to the present disclosure.

[0012] Figure 9 This is a schematic diagram of a fan blade assembly according to the present disclosure. Detailed Implementation

[0013] Reference will now be made in detail to the present embodiments of this disclosure, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerals and letter reference numerals to designate features in the drawings. Similar or analogous reference numerals in the drawings and description have been used to designate similar or analogous portions of this disclosure.

[0014] The term "exemplary" is used herein to mean "used as an example, instance, or illustration." Any implementation described herein as "exemplary" is not necessarily to be construed as superior or better than other implementations. Furthermore, unless explicitly stated otherwise, all embodiments described herein should be considered exemplary.

[0015] For the purposes described below, the terms “upper,” “lower,” “right,” “left,” “vertical,” “horizontal,” “top,” “bottom,” “lateral,” “longitudinal,” and their derivatives should be associated with this disclosure as oriented as it is in the accompanying drawings. However, it should be understood that various alternative variations may be assumed in this disclosure unless explicitly stated otherwise. It should also be understood that the specific devices shown in the drawings and described in the following description are merely exemplary embodiments of this disclosure. Therefore, the specific dimensions and other physical characteristics associated with the embodiments disclosed herein should not be considered limiting.

[0016] As used herein, the terms “first,” “second,” and “third” are used interchangeably to distinguish one component from another and are not intended to indicate the location or importance of the individual components.

[0017] The terms "front" and "rear" refer to relative positions within a gas turbine engine or carrier, and specifically to the normal operating posture of the gas turbine engine or carrier. For example, in the case of a gas turbine engine, "front" refers to the position closer to the engine inlet, while "rear" refers to the position closer to the engine nozzle or exhaust port.

[0018] The terms "upstream" and "downstream" refer to the relative directions of fluid flow within a fluid path. For example, "upstream" refers to the direction from which the fluid flows, and "downstream" refers to the direction from which the fluid flows.

[0019] Unless otherwise stated herein, the terms “connection,” “fixed,” “attached to,” etc., refer to both direct connection, fixation, or attachment, and indirect connection, fixation, or attachment via one or more intermediate components or features.

[0020] Unless the context clearly indicates otherwise, the singular forms “a,” “a,” and “the” include plural references.

[0021] As used throughout the specification and claims, approximate language is applied to modify any quantitative expression that may allow for variation without altering its underlying function. Therefore, values ​​modified by terms such as “about,” “approximately,” and “substantially” are not limited to specified precise values. In at least some cases, approximate language may correspond to the precision of the instrument used to measure the value, or the precision of the method or machine used to construct or manufacture the component and / or system. For example, approximate language may refer to margins of 1%, 2%, 4%, 10%, 15%, or 20%. These approximate margins may be applied to a single value, to either end of a range defining a numerical value, or to margins between two ends, and / or between the ends.

[0022] Throughout this specification and claims, scope limitations are combined and interchanged, and unless the context or language otherwise indicates otherwise, such scopes are identified and include all subscopes contained herein. For example, all scopes disclosed herein include endpoints, and endpoints may be combined independently of each other.

[0023] The term “turbine” or “turbomachinery” refers to a machine that includes one or more compressors, a heating section (e.g., a combustion section), and one or more turbines that together generate torque output.

[0024] The term "gas turbine engine" refers to an engine that has a turbine as its power source, in whole or in part. Examples of gas turbine engines include turbofan engines, turboprop engines, turbojet engines, turboshaft engines, etc.

[0025] The term "combustion section" refers to any heat addition system used in a turbine. For example, the term "combustion section" can refer to a section that includes one or more of a knock combustion assembly, a rotary detonation combustion assembly, a pulse detonation combustion assembly, or other suitable heat addition assemblies. In some exemplary embodiments, the combustion section may include an annular burner, a can burner, a tubular burner, a vortex burner (TVC), or other suitable combustion systems, or combinations thereof.

[0026] When used with compressors, turbines, shafts, or spool components, unless otherwise specified, the terms “low” and “high,” or their respective comparatives (e.g., “lower” and “higher,” where applicable), refer to relative speeds within the engine. For example, “low-speed turbine” or “low-turbine” defines a component configured to operate at a rotational speed (e.g., the maximum permissible rotational speed) lower than that of a “high-speed turbine” or “high-speed turbine” at the engine.

[0027] This disclosure generally relates to fan blade assemblies. More specifically, this disclosure relates to fan blade assemblies including a midspan shroud, and gas turbine engines including the fan blade assembly.

[0028] The midspan shroud separates the interior portion of the fan assembly from the exterior portion. For example, the midspan plane defined by the midspan shroud may be orthogonal to the inner plane defined by the inner portion and the outer plane defined by the outer portion. Additionally or alternatively, the midspan plane may be aligned along the circumferential direction of the gas turbine engine and may be orthogonal to the radial direction of the gas turbine engine.

[0029] A gas turbine engine includes a turbine. The turbine includes, for example, a casing defining a first upstream end at a first inlet in the core airflow path through the turbine. The casing at least partially separates a bypass airflow passage from the core airflow path.

[0030] The midspan shield may be aligned with the first upstream end of the housing. Specifically, the midspan shield may at least partially define the bypass airflow passage and the core airflow path. The engine may include multiple fan blade assemblies, each including a corresponding midspan shield. The multiple midspan shields may define a cylindrical shape defining a second inlet and a second upstream end, and extending upstream of the bypass airflow passage and the core airflow passage.

[0031] The second inlet is upstream of the first inlet, and the second upstream end is upstream of the first upstream end, to extend the upstream inlet location of the bypass airflow passage and the core airflow passage. For example, the second upstream end is upstream of the fan. Multiple mid-span shrouds are used to separate the bypass airflow passage and the core airflow path.

[0032] The fan blade assembly may be part of a fan, and further includes at least one actuator operatively coupled to an inner portion, an outer portion, or both of the fan blade assembly to control the pitch of the inner portion, the pitch of the outer portion, or both. In some configurations, the at least one actuator may be configured to change the pitch of the outer portion relative to the pitch of the inner portion during various operating conditions of the gas turbine engine.

[0033] For example, during reverse thrust operation, at least one actuator can be configured to move the outer portion to the reverse thrust pitch while the inner portion is at the forward thrust pitch. This allows the fan blade assembly to continue supplying the desired airflow to the turbine of the gas turbine engine during reverse thrust operation (e.g., to ensure it does not stall), while still generating reverse thrust for the gas turbine engine.

[0034] Referring now to the accompanying drawings, where the same numbers indicate the same elements throughout all the drawings. Figure 1 This is a schematic cross-sectional view of a gas turbine engine according to an exemplary embodiment of the present disclosure. More specifically, for Figure 1In one embodiment, the gas turbine engine is an aircraft turbofan engine 10, which is configured to be mounted on the aircraft, for example, in an underwing configuration or a tail-mount configuration.

[0035] like Figure 1 As shown, the turbofan engine 10 defines an axial direction A (extending parallel to the longitudinal centerline 12 provided for reference), a radial direction R, and a circumferential direction C (i.e., the direction extending about the axial direction A; see...). Figure 2 ).

[0036] Typically, the turbofan engine 10 includes a fan section 14 and a turbine 16 disposed downstream of the fan section 14. The turbine 16 is sometimes also referred to, or alternatively referred to, as the "core turbine engine".

[0037] An exemplary turbine 16 includes a housing 18 defining a first inlet 20. The housing 18 may be substantially tubular and the first inlet 20 may be annular. The housing 18 surrounds, in a series flow relationship: a compressor section including a first boost or low-pressure (LP) compressor 22 and a second high-pressure (HP) compressor 24; a combustion section including a combustor 26; a turbine section including a first high-pressure (HP) turbine 28 and a second low-pressure (LP) turbine 30; and an exhaust nozzle section 32.

[0038] A high-pressure (HP) shaft 34 or spool drives the HP turbine 28 to the HP compressor 24. A low-pressure (LP) shaft 36 or spool drives the LP turbine 30 to the LP compressor 22. The compressor section, combustion section, turbine section, and injection exhaust nozzle section 32 are arranged in a sequential flow order and together define the core airflow path 37 through the turbine 16.

[0039] Fan section 14 includes a variable pitch single-stage fan 38. Turbine 16 is operatively coupled to and drives fan 38.

[0040] The fan 38 includes a plurality of fan blade assemblies 40 spaced apart from the disk 42. As shown, the fan blade assemblies 40 extend outward from the disk 42 in a generally radial direction R.

[0041] As described in further detail below, the fan blade assembly 40 is operably coupled to one or more suitable actuators 44 and is rotatable.

[0042] The fan blade assembly 40, disk 42, and actuator 44 can rotate together about the longitudinal centerline 12 via a power gearbox 46 across an LP shaft 36. The power gearbox 46 includes multiple gears for reducing the rotational speed of the LP shaft 36 to a more efficient fan rotation speed. Thus, in the depicted embodiment, the turbine 16 is operatively coupled to the fan 38 via the power gearbox 46.

[0043] During operation of the turbofan engine 10, fan 38 defines the fan pressure ratio. As used herein, the term "fan pressure ratio" refers to the ratio of the air pressure immediately downstream of fan 38 to the air pressure immediately upstream of fan 38.

[0044] The disk 42 is covered by a rotatable forward nacelle 48, which is aerodynamically shaped to facilitate airflow through the plurality of fan blade assemblies 40. Additionally, the exemplary fan section 14 may include an annular fan housing or outer nacelle 50, which at least partially and, in the depicted embodiment, circumferentially surrounds at least a portion of the fan 38 and turbine 16. In other embodiments, the turbofan engine is an open rotor engine and the outer nacelle 50 is omitted.

[0045] Furthermore, in the depicted embodiment, the outer nacelle 50 is supported relative to the turbine 16 by a plurality of circumferentially spaced outlet guide vanes 52. A downstream section 54 of the outer nacelle 50 extends over the outer portion of the turbine 16 to define a bypass airflow passage 56 therebetween.

[0046] During operation of the turbofan engine 10, a certain amount of air 58 enters the turbofan engine 10 through the external nacelle 50 and / or the relevant inlet 60 of the fan section 14. As the certain amount of air 58 passes through the fan blade assembly 40, a first portion of the air 58 (bypass airflow 62) is directed or directed into the bypass airflow passage 56, while a second portion of the air 58 (core airflow 64) is directed or directed into the core airflow path 37.

[0047] As the core airflow 64 is directed through the LP compressor 22 and HP compressor 24 and into the combustor 26, the pressure of the core airflow 64 increases. More specifically, the compressor section including the LP compressor 22 and HP compressor 24 defines the total pressure ratio during operation of the turbofan engine 10 at rated speed. The total pressure ratio refers to the ratio of the outlet pressure of the compressor section (i.e., the pressure of the core airflow 64 at the rear end of the compressor section) to the inlet pressure of the compressor section (i.e., the pressure of the core airflow 64 at the first inlet 20 of the compressor section).

[0048] The compressed core gas flow 64 from the compressor section is mixed with fuel and burned in the combustion section to provide combustion gas 66. The combustion gas 66 is directed from the combustor 26 through the HP turbine 28, where a first portion of the thermal and / or kinetic energy from the combustion gas 66 is extracted via a successive stage of HP turbine stator blades 68 coupled to the housing 18 and a plurality of HP turbine rotor blades 70 coupled to the HP shaft 34, thus rotating the HP shaft 34 to support the operation of the HP compressor 24.

[0049] Combustion gas 66 is then directed through LP turbine 30, where a second portion of thermal and / or kinetic energy is extracted from combustion gas 66 via a successive stage of LP turbine stator blades 72 coupled to housing 18 and a plurality of LP turbine rotor blades 74 coupled to LP shaft 36, thereby causing LP shaft 36 to rotate, thereby supporting the operation of LP compressor 22 and / or rotation of fan 38.

[0050] Combustion gas 66 is then directed through the injector exhaust nozzle section 32 of turbine 16 to provide propulsive thrust. Simultaneously, the pressure of bypass airflow 62 increases significantly as it is directed through bypass airflow passage 56 before exiting from the fan nozzle exhaust section 76 of turbofan engine 10, also providing propulsive thrust. HP turbine 28, LP turbine 30, and injector exhaust nozzle section 32 at least partially define a hot gas path 78 for directing combustion gas 66 through turbine 16.

[0051] It should be understood that Figure 1 The exemplary turbofan engine 10 depicted is merely an example, and in other exemplary embodiments, the turbofan engine 10 may have any other suitable configuration. The aspects of this disclosure can be used with any other suitable aero gas turbine engine (e.g., turboshaft engine, turboprop engine, turbojet engine, etc.). Furthermore, the aspects of this disclosure can also be used with any other land-based gas turbine engine, such as any aero-derivative gas turbine engine, such as a marine gas turbine engine.

[0052] Other gas turbine engines to which this disclosure can be applied may have alternative configurations. As an example, such an engine may have an alternative number of interconnecting shafts (e.g., two) and / or an alternative number of compressors and / or turbines. Furthermore, the engine may include a gearbox disposed in the transmission system from the turbine to the compressor and / or fan.

[0053] It should be understood that Figure 1 The exemplary turbofan engine 10 depicted is merely an example, and in other exemplary embodiments, the turbofan engine 10 may have any other suitable configuration. For example, aspects of this disclosure can be used with any other suitable aviation gas turbine engine (e.g., turboshaft engine, turboprop engine, turbojet engine, etc.). Furthermore, aspects of this disclosure can also be used with any aero-derivative gas turbine engine, such as a marine gas turbine engine.

[0054] Other gas turbine engines to which this disclosure can be applied may have alternative configurations. For example, such engines may have an alternative number of interconnecting shafts (e.g., two) and / or an alternative number of compressors and / or turbines. Furthermore, the engine may not include a gearbox disposed in the drivetrain from the turbine to the compressor and / or fan, and may be configured as a non-pipeline gas turbine engine (e.g., excluding the outer nacelle 50), etc.

[0055] Still referencing Figure 2 , Figure 2 Provided Figure 1 A schematic front view of the fan section 14 shows the turbine 16 housing 18 with a first upstream end 80 positioned at a radial distance 82 from the longitudinal centerline 12. The first upstream end 80 defines the separation point between the bypass airflow passage 56 and the core airflow path 37. The housing 18 at least partially defines and directs the bypass airflow 62 (e.g., a first portion of air 58) into the bypass airflow passage 56, and defines and directs the core airflow 64 (e.g., a second portion of air 58) into the core airflow path 37.

[0056] Each fan blade assembly 40 includes an inner portion 90 in the radial direction R, a mid-span shroud 92, and an outer portion 94 in the radial direction R. The mid-span shroud 92 is aligned with a first upstream end 80 of the housing 18 (e.g., the radial distance from the longitudinal centerline 12 to the mid-span shroud 92 is equal to the radial distance 82).

[0057] refer to Figure 2 It should be understood that fan section 14 includes a plurality of fan blade assemblies 40, each of which includes a midspan shroud 92. The midspan shrouds 92 of the fan blade assemblies 40 together have a cylindrical shape extending from a first upstream end 80 of the first inlet 20 (e.g., the cylindrical shape includes the first upstream end 80 and a second upstream end 84 of the second inlet 86 upstream of the first inlet 20 (e.g., the upstream end of the midspan shroud 92)), extending from the upstream end of the bypass airflow passage 56 and / or at least partially defining the bypass airflow passage 56, and extending from the upstream end of the core airflow path 37 and / or at least partially defining the core airflow path 37. Alternatively, the midspan shroud 92 may separate air 58 entering the inlet 60 at or upstream of the fan 38 to isolate the bypass airflow 62 and the core airflow 64 before they flow into the bypass airflow passage 56 and the core airflow path 37, respectively.

[0058] The mid-span shield 92 separates the inner portion 90 from the outer portion 94. (Reference) Figure 2 The length or distance from the root 96 of the fan blade assembly 40 to the tip 98 of the fan blade assembly 40 can be referred to as the span of the fan blade assembly 40. (Reference) Figure 1The width or distance between the edges 100 and 102 of the fan blade assembly 40 can be referred to as the chord of the fan blade assembly 40.

[0059] The midspan shroud 92 can be positioned along the span of the fan blade assembly 40 closer to the root 96 than to the tip 98. The midspan plane of the midspan shroud 92 can be substantially orthogonal to the inner plane of the inner portion 90 and the outer plane of the outer portion 94 (not marked).

[0060] As described above with respect to fan blade assembly 40, the inner portion 90 and / or the outer portion 94 in the radial direction R are operatively coupled to one or more actuators 44. For example, the actuators 44 may be configured to (commonly or independently) change the pitch of the inner portion 90 and / or the outer portion 94 (e.g., rotate about the pitch axis P).

[0061] refer to Figure 3 , Figure 5 and Figure 7 Each figure provides an exemplary aspect according to this disclosure. Figure 2 An end view of the fan blade assembly 40 of fan section 14, wherein the outer pitch 114 of the outer portion 94 can vary relative to the inner pitch 110 of the inner portion 90. For example, each of the inner pitch 110 and the outer pitch 114 is variable. As another example, the inner pitch 110 can be fixed, while the outer pitch 114 can be variable. (See also: Special Reference) Figure 3 The outer pitch 114 and the inner pitch 110 are shown with different positive pitch angles.

[0062] Because the bypass airflow 62 and the core airflow 64 can be adjusted independently, the bypass ratio can be adjusted. The ratio between the amount of bypass airflow 62 through the bypass airflow channel 56 (i.e., the first portion of air 58) and the amount of core airflow 64 through the core airflow path 37 (i.e., the second portion of air 58) is called the bypass ratio. For example, now specifically refer to... Figure 3 and Figure 5 , Figure 3 The structure can have the same Figure 5 The different bypass ratios are due to the different pitches 110 and 114 of the radial inner portion 90 and the radial outer portion 94. Figure 3 The pitches are different, while those with pitches of 110 and 114 are... Figure 5 Same as above.

[0063] Now for special reference Figure 5 and Figure 7 And still referencing Figure 4 and Figure 6 Each figure provides an exemplary aspect according to this disclosure. Figure 2A perspective view of the fan blade assembly 40 of fan section 14. According to the exemplary method, actuator 44 can move the outer pitch 114 of the outer portion 94 to a reverse pitch angle while maintaining the inner pitch 110 of the inner portion 90 at a positive pitch angle. Figure 4-5 In this context, the outer pitch 114 and the inner pitch 110 are the same or similar, and each represents a positive pitch angle (e.g., positive thrust operation mode). Figure 5 The image shows an end view of the fan blade assembly 40 as seen from the tip 98. Figure 4 This is a 3D view of the fan blade assembly 40.

[0064] refer to Figure 6-7 The outer portion 94 is rotated such that the outer pitch 114 is at a reverse pitch angle, while the inner pitch 110 of the inner portion 90 remains at a positive pitch angle. In this configuration (e.g., reverse thrust operation mode), the inner portion 90 is maintained at a positive pitch to keep the core airflow 64 in the positive direction when the bypass airflow 62 is in the reverse direction due to the reverse pitch of the outer portion 94.

[0065] This minimizes the problem of lack of core airflow 64 in turbine 16 during reverse thrust operation. The inner section 90 maintains some flow in the positive direction to the core airflow path 37, and because the mid-span shield 92 separates the core airflow 64 from the bypass airflow 62, there is less interference between the airflows 62 and 64 when moving in different directions.

[0066] It should be understood that, as used herein, the term "positive pitch angle" for a rotor blade refers to the pitch angle of the rotor blade, whereby the rotor blade is configured to provide airflow from a forward position to a rearward position to generate, for example, positive thrust. In contrast, as used herein, the term "reverse pitch angle" for a rotor blade refers to the pitch angle of the rotor blade, whereby the rotor blade is configured to provide airflow from a rearward position to a forward position to generate, for example, reverse thrust.

[0067] refer to Figure 8 A schematic diagram of a fan blade assembly 40 according to the present disclosure is provided. The fan blade assembly 40 includes a first actuator 140 operably coupled to an inner portion 90 and a second actuator 144 operably coupled to an outer portion 94 of the fan blades. Here, a mid-span shroud 92 is fixed in place (e.g., fixed to a disc 42), and the inner portion 90 and the outer portion 94 are rotatable relative to the disc 42 or other structures of the engine.

[0068] For example, bearings 150 and 154 are configured to allow rotation of the inner portion 90 and the outer portion 94. A first actuator 140 controls the pitch of the inner portion 90 about the pitch axis P, and a second actuator 144 controls the pitch of the outer portion 94 about the pitch axis P.

[0069] refer to Figure 9 , Figure 9 A schematic diagram of a fan blade assembly 40 according to this disclosure is provided, the fan blade assembly 40 including a second actuator 144 operatively coupled to an outer portion 94 of a fan blade. Here, the inner portion 90 and the mid-span shroud 92 are fixed in place (e.g., fixed to a disc 42). The outer portion 94 is rotatable relative to the disc 42 or other structures of the engine. For example, a bearing 154 is configured to allow rotation of the outer portion 94. The second actuator 144 controls the pitch of the outer portion 94 about a pitch axis P.

[0070] The controller 160 is configured to control one or more of the actuators 140, 144 to change the pitch or pitch position of the outer portion 94 relative to the pitch of the inner portion 90 during various operating conditions of the turbofan engine 10. For example, under reverse thrust operation, the controller 160 is configured to control a second actuator 144 to move the outer portion 94 to a reverse pitch position (e.g., a pitch that generates reverse thrust) while the inner portion 90 is in a forward pitch position (e.g., a pitch that generates forward thrust). To be in the forward pitch position, the inner portion 90 may have a fixed forward pitch, or the controller 160 may control a first actuator 140 to move the inner portion 90 to the forward pitch position.

[0071] This configuration allows the fan blade assembly 40 to continue supplying the desired airflow to the turbine 16 of the turbofan engine 10 during reverse thrust operation (e.g., to ensure it does not stall), while still generating reverse thrust for the turbofan engine 10.

[0072] The controller 160 can control the actuators 140 and 144 to independently control the pitch of the inner portion 90 and the outer portion 94 to have a positive pitch and generate positive thrust or airflow, or to have a negative pitch and generate negative thrust or airflow.

[0073] Typically, the exemplary controller 160 is configured to receive data or commands (e.g., mode commands) from one or more systems and, for example, make control decisions based on the received data.

[0074] In one or more exemplary embodiments, controller 160 may be a standalone controller, or alternatively, it may be integrated into one or more of a controller for turbofan engine 10, a controller for an aircraft including turbofan engine 10, etc.

[0075] With particular reference to the operation of controller 160, in at least some embodiments, controller 160 may include one or more computing devices 164. The one or more computing devices 164 may include one or more processors 164A and one or more memory devices 164B.

[0076] One or more processors 164A may include any suitable processing device, such as a microprocessor, microcontroller, integrated circuit, logic device and / or other suitable processing device.

[0077] One or more memory devices 164B may include one or more computer-readable media, including but not limited to non-transitory computer-readable media, RAM, ROM, hard disk drives, flash drives, and / or other memory devices. One or more memory devices 164B may store information accessible by one or more processors 164A, including computer-readable instructions 164C executable by one or more processors 164A. Instructions 164C may be any set of instructions that, when executed by one or more processors 164A, cause one or more processors 164A to operate. In some embodiments, instructions 164C may be executed by one or more processors 164A to cause one or more processors 164A to operate, for example, any operations and functions configured for the controller 160 and / or one or more computing devices 164, operations (e.g., modes or methods) for operating the turbofan engine 10 as described herein, and / or any other operations or functions of one or more computing devices 164.

[0078] Instruction 164C can be software written in any suitable programming language or can be implemented in hardware. Additionally and / or alternatively, instruction 164C can be executed in logically and / or virtually decoupled threads on one or more processors 164A.

[0079] One or more memory devices 164B may further store data 164D that can be accessed by one or more processors 164A. For example, data 164D may include data indicating power flow, data indicating engine / aircraft operating conditions, and / or any other data and / or information described herein.

[0080] One or more computing devices 164 may also include a network interface 164E for communicating, for example, with other components of the turbofan engine 10, an aircraft incorporating a gas turbine engine, etc. For example, in the depicted embodiment, the turbofan engine 10 may operate in several operating modes. The controller 160 is operatively coupled to one or more aircraft systems (e.g., a flight management system or other aircraft control system) via, for example, a network interface, such that the controller 160 can receive data or commands indicating various modes.

[0081] The network interface 164E may include any suitable components for use with one or more network interfaces, including, for example, a transmitter, receiver, port, controller, antenna and / or other suitable components.

[0082] The techniques discussed herein refer to computer-based systems, actions taken by computer-based systems, information sent to computer-based systems, and information received from computer-based systems. Those skilled in the art will recognize that the inherent flexibility of computer-based systems allows for a wide variety of possible configurations, combinations, and divisions of tasks and functions between and within components. For example, the processes discussed herein can be implemented using a single computing device or multiple computing devices working in combination. Databases, memories, instructions, and applications can be implemented on a single system or distributed across multiple systems. Distributed components can operate sequentially or in parallel.

[0083] This written description uses examples to disclose this disclosure, including best practices, and to enable any person skilled in the art to practice this disclosure, including making and using any device or system and methods of making any combination. The patent scope of this disclosure is defined by the claims, but may include other examples that would occur to a person skilled in the art. Such other examples are intended to fall within the scope of the claims if they include structural elements that are not indistinguishable from the literal language of the claims, or if they include equivalent structural elements that are not substantially different from the literal language of the claims.

[0084] Further details are provided by the following topics:

[0085] A gas turbine engine defining a radial direction includes: a fan, the fan including a fan blade assembly, the fan blade assembly including: an inner portion along the radial direction; a mid-span shroud; and an outer portion along the radial direction; and at least one actuator operatively coupled to the inner portion, the outer portion, or both, to control the inner pitch of the inner portion, the outer pitch of the outer portion, or both.

[0086] According to one or more of these clauses, a gas turbine engine is wherein the at least one actuator is operatively coupled to the outer portion, wherein the inner portion has a fixed positive pitch, and wherein the outer portion has a variable pitch.

[0087] According to one or more of these clauses, the gas turbine engine wherein the outer portion is movable between a forward pitch position and a reverse pitch position by the at least one actuator.

[0088] According to one or more of these clauses, in a gas turbine engine, the at least one actuator is configured to change the outer pitch of the outer portion relative to the inner pitch of the inner portion.

[0089] According to one or more of these clauses, each of the inner portion and the outer portion has a variable pitch.

[0090] According to one or more of these clauses, the gas turbine engine wherein the midspan shield separates the inner portion from the outer portion.

[0091] According to one or more of these clauses, in a gas turbine engine, the midspan shroud is closer to the root of the fan blade assembly than to the tip of the fan blade assembly.

[0092] According to one or more of these clauses, the gas turbine engine has a mid-span plane that is orthogonal to the inner plane of the inner portion and the outer plane of the outer portion.

[0093] According to one or more of these clauses, the gas turbine engine wherein the mid-span shroud at least partially defines at least one of the bypass airflow passage and the core airflow path.

[0094] The gas turbine engine according to one or more of these clauses further includes: a turbine drivenly connected to the fan, the turbine including a housing defining a first upstream end and a first inlet of a core airflow path, wherein a bypass airflow passage and the core airflow path are separated at least by the housing.

[0095] According to one or more of these clauses, the gas turbine engine wherein the mid-span shroud is aligned with the first upstream end.

[0096] According to one or more of these clauses, the gas turbine engine is wherein the first upstream end is located at a first radial distance from the longitudinal centerline of the gas turbine engine, wherein the midspan shroud is located at a second radial distance from the longitudinal centerline of the gas turbine engine, and wherein the first radial distance is substantially equal to the second radial distance.

[0097] According to one or more of these clauses, the gas turbine engine, wherein the fan includes a plurality of fan blade assemblies, each of the plurality of fan blade assemblies including a corresponding midspan shroud, and wherein the midspan shrouds of the plurality of fan blade assemblies together define a cylindrical shape, the cylindrical shape including a second inlet upstream of the first inlet of the core airflow path.

[0098] According to one or more of these clauses, the gas turbine engine wherein the mid-span shield separates the bypass airflow passage from the core airflow path.

[0099] A fan blade assembly for defining a radial direction of a gas turbine engine, the fan blade assembly comprising: an inner portion along the radial direction of the gas turbine engine, the inner portion defining an inner pitch when mounted in the gas turbine engine; an outer portion along the radial direction of the gas turbine engine, the outer portion defining an outer pitch when mounted in the gas turbine engine, wherein the outer portion is rotatable relative to the inner portion such that the outer pitch is variable relative to the inner pitch; and a mid-span shroud positioned between the inner portion and the outer portion.

[0100] According to one or more of these clauses, the fan blade assembly has an inner portion having a fixed positive pitch and an outer portion having a variable pitch.

[0101] According to one or more of these clauses, the fan blade assembly wherein the midspan shroud separates the inner portion from the outer portion.

[0102] According to one or more of these clauses, the fan blade assembly is located closer to the root of the fan blade assembly than to the tip of the fan blade assembly.

[0103] According to one or more of these clauses, the fan blade assembly wherein the mid-span plane of the mid-span shroud is orthogonal to the inner plane of the inner portion and the outer plane of the outer portion.

[0104] A method of operating a gas turbine engine having a fan with fan blades, the method comprising: operating the gas turbine engine in a forward thrust operating mode; and operating the gas turbine engine in a reverse thrust operating mode, wherein operating the gas turbine engine in the reverse thrust operating mode comprises changing the outer pitch of an outer portion of the fan blade relative to the inner pitch of an inner portion of the fan blade, wherein a mid-span shroud separates the inner portion from the outer portion.

Claims

1. A gas turbine engine with a defined radial direction, characterized in that, include: A fan, the fan including a plurality of fan blade assemblies, each of the plurality of fan blade assemblies including: The inner portion along the radial direction; Mid-span shield; and The outer portion along the radial direction; and At least one actuator, operatively coupled to the inner portion, the outer portion, or both, to control the inner pitch of the inner portion, the outer pitch of the outer portion, or both; A turbine, droneally coupled to the fan, includes a housing defining a first upstream end and a first inlet to a core airflow path, wherein a bypass airflow passage and the core airflow path are separated at least by the housing. The mid-span shroud of the plurality of fan blade assemblies together includes a second inlet located upstream of the first inlet in the core airflow path; Wherein, the first upstream end and the mid-span shield define the separation between the bypass airflow channel and the core airflow path; Each of the inner portion and the outer portion has a variable pitch.

2. The gas turbine engine according to claim 1, characterized in that, in, The at least one actuator is operatively coupled to the outer portion.

3. The gas turbine engine according to claim 2, characterized in that, in, The outer portion can be moved between a forward pitch position and a reverse pitch position by the at least one actuator.

4. The gas turbine engine according to claim 1, characterized in that, in, The at least one actuator is configured to change the outer pitch of the outer portion relative to the inner pitch of the inner portion.

5. The gas turbine engine according to claim 1, characterized in that, in, The mid-span shield separates the inner portion from the outer portion.

6. The gas turbine engine according to claim 1, characterized in that, in, The midspan shroud is closer to the root of the fan blade assembly than to the tip of the fan blade assembly.

7. The gas turbine engine according to claim 1, characterized in that, in, The mid-span plane of the mid-span shield is orthogonal to the inner plane of the inner part and the outer plane of the outer part.

8. The gas turbine engine according to claim 1, characterized in that, in, The mid-span protective cover is aligned with the first upstream end.

9. The gas turbine engine according to claim 1, characterized in that, in, The first upstream end is positioned at a first radial distance from the longitudinal centerline of the gas turbine engine, wherein the mid-span shroud is positioned at a second radial distance from the longitudinal centerline of the gas turbine engine, and wherein the first radial distance is equal to the second radial distance.

10. The gas turbine engine according to claim 1, characterized in that, in, Each of the midspan shrouds of the plurality of fan blade assemblies includes first and second opposing edges spaced apart from each other in the circumferential direction on opposite sides of the pitch axis of the respective fan blade assembly.

11. The gas turbine engine according to claim 10, characterized in that, in, The mid-span shield separates the bypass airflow channel from the core airflow path.

12. A fan blade assembly for a gas turbine engine defining a radial direction, characterized in that, The fan blade assembly includes: An inner portion along the radial direction of the gas turbine engine, when installed in the gas turbine engine, defines an inner pitch; An outer portion along the radial direction of the gas turbine engine, when installed in the gas turbine engine, defines an outer pitch, wherein the outer portion is rotatable relative to the inner portion, such that the outer pitch is variable relative to the inner pitch; and A midspan shroud is positioned between the inner portion and the outer portion, wherein the midspan shroud includes first and second opposing edges spaced apart from each other in the circumferential direction on opposite sides of the pitch axis of the fan blade assembly, wherein the midspan shroud separates the inner portion from the outer portion, wherein each of the inner portion and the outer portion has a variable pitch.

13. The fan blade assembly according to claim 12, characterized in that, in, The midspan shroud is closer to the root of the fan blade assembly than to the tip of the fan blade assembly.

14. The fan blade assembly according to claim 12, characterized in that, in, The mid-span plane of the mid-span shield is orthogonal to the inner plane of the inner part and the outer plane of the outer part.

15. A method of operating a gas turbine engine having a fan with a fan blade assembly, characterized in that, The method includes: Operating the gas turbine engine in forward thrust mode; and Operating the gas turbine engine in a reverse thrust operation mode includes changing the outer pitch of the outer portion of the fan blade assembly relative to the inner pitch of the inner portion of the fan blade assembly, wherein a midspan shroud separates the inner portion from the outer portion, wherein the midspan shroud includes first and second opposing edges spaced apart from each other in the circumferential direction on opposite sides of the pitch axis of the fan blade assembly, wherein each of the inner portion and the outer portion has a variable pitch.

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