Partial span shield for pitch control aircraft

Abstract

This document discloses a partial span shield for a pitch control aircraft. An example device includes: a first portion extending from a first airfoil, the first portion including a first surface that is convex; a second portion extending from a second airfoil toward the first airfoil, the second portion including a second surface circumferentially adjacent to the first airfoil, the second surface being concave; and an interface formed by the first and second surfaces, the interface reacting to a circumferential load between the first and second airfoils.

Description

Technical Field

[0001] This disclosure generally relates to turbines, and more specifically, to a partial span shield for pitch control aircraft. Background Technology

[0002] Gas turbine engines typically consist of an inlet section, compressor section, combustion section, turbine section, and exhaust section in a sequential flow order. During operation, air enters the inlet section and flows to the compressor section, where one or more axial compressors progressively compress the air until it reaches the combustion section, producing combustion gases. The combustion gases flow from the combustion section through a hot gas path defined within the turbine section, and then exit the turbine section via the exhaust section.

[0003] Pitch variations allow rotating airfoils to rotate along their radial axes to change blade pitch. This pitch control enables these airfoils to maintain optimal angle of attack under various environmental and aircraft conditions. Pitch control on rotating airfoils improves the overall efficiency of gas turbine engines. Attached Figure Description

[0004] Figure 1 This is a cross-sectional view of an example turbofan gas turbine engine in which the examples disclosed herein can be implemented.

[0005] Figure 2 This is a cross-sectional view of an example open rotary engine in which the examples disclosed herein can be implemented.

[0006] Figure 3 It includes a partial span shield. Figure 1 Front view of an open rotor.

[0007] Figure 4 yes Figure 3 A top view of part of the span of the protective shield.

[0008] Figure 5 This is a top view of the replacement portion of the shield.

[0009] The accompanying drawings are not drawn to scale. Instead, the thickness of layers or regions may be enlarged in the drawings. Although layers and regions with sharp lines and boundaries are shown in the drawings, some or all of these lines and / or boundaries may be idealized. In reality, boundaries and / or lines may be unobservable, mixed, and / or irregular. Generally, the same reference numerals will be used throughout the accompanying drawings and accompanying written description to refer to the same or similar parts. As used herein, unless otherwise stated, the term "above" describes the relationship of two parts relative to the Earth. If the second part has at least one section between the Earth and the first part, then the first part is above the second part. Similarly, as used herein, the first part is "below" the second part when the first part is closer to the Earth than the second part. As stated above, the first part may be above or below the second part by one or more of the following: there are other parts in between, there are no other parts in between, the first part and the second part are in contact, or the first part and the second part are not in direct contact with each other. As used herein, a statement that any part (e.g., layer, film, region, area, or plate) is on another part in any manner (e.g., positioned, located, set, or formed, etc.) indicates that the referenced part is either in contact with or above the other part, with one or more intermediate parts located between them. As used herein, unless otherwise stated, a connecting reference (e.g., attachment, coupling, connection, and joining) may include intermediate members between elements referenced by the connecting reference and / or relative movement between those elements. Therefore, a connecting reference does not necessarily imply that two elements are directly connected and have a fixed relationship with each other. As used herein, a statement that any part is “in contact” with another part is defined to mean that there is no intermediate part between the two parts.

[0010] Unless otherwise specifically stated, descriptors used herein (such as "first," "second," "third," etc.) do not assign or otherwise indicate any meaning of priority, physical order, arrangement in a list, and / or any sorting, but are merely used as labels and / or arbitrary names to distinguish elements for ease of understanding the disclosed examples. In some examples, the descriptor "first" may be used to refer to an element in the detailed description, while different descriptors (e.g., "second" or "third") may be used in the claims to refer to the same element. In such cases, it should be understood that such descriptors are only used to clearly identify those elements that may, for example, otherwise share the same name. As used herein, "approximately" and "about" refer to dimensions that may be inaccurate due to manufacturing tolerances and / or other real-world defects. Detailed Implementation

[0011] Aircraft include engines that act as propulsion systems to generate mechanical power and force (such as thrust). A gas turbine (also known as a combustion turbine or turbine engine) is an internal combustion engine that can be implemented in the propulsion system of an aircraft. For example, a gas turbine can be implemented in combination with a turbofan engine or a turbojet engine. Gas turbines also have important applications in fields such as industrial power generation.

[0012] In the following detailed description, reference is made to the accompanying drawings, which form a part thereof, in which specific examples that can be practiced are illustrated by way of illustration. These examples are described in sufficient detail to enable those skilled in the art to practice the subject matter, and it should be understood that other examples may be used. Therefore, the following detailed description is provided to describe exemplary embodiments and is not to be construed as limiting the scope of the subject matter described in this disclosure. Certain features from different aspects of the following description may be combined to form new aspects of the subject matter discussed below.

[0013] In describing the elements of the various embodiments of this disclosure, the articles “a,” “an,” “the,” and “the” are intended to indicate the presence of one or more elements. The terms “first,” “second,” etc., do not indicate any order, quantity, or importance, but are used to distinguish one element from another. The terms “comprising,” “including,” and “having” are intended to include and mean that additional elements may be present in addition to those listed. As used herein, the terms “connected to,” “linked to,” etc., indicate that an object (e.g., a material, element, structure, component, etc.) may be connected to or linked to another object, regardless of whether the object is directly connected to or linked to another object, or whether one or more intermediate objects exist between the object and the other.

[0014] 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.

[0015] In some of the examples described herein, the location of features on the airfoil is described as a percentage of the blade's span. In such examples disclosed herein, the percentage refers to the feature's spanwise location relative to the blade root. Specifically, a feature at 0% span is located at the root of the airfoil, a feature at 100% span is located at the tip of the airfoil, a feature at 50% span is located between the tip and the root, and so on.

[0016] In some examples used in this article, the term "substantially" is used to describe the relationship between two parts that is within three degrees of the relationship (e.g., substantially collinear relationship within three degrees of linearity, substantially perpendicular relationship within three degrees of verticality, substantially parallel relationship within three degrees of parallelism, etc.).

[0017] As used herein, the terms “axial” and “longitudinal” refer to directions parallel to the centerline axis of the gas turbine (e.g., turbofan engine, core gas turbine engine, etc.), while “radial” refers to a direction perpendicular to the axial direction, and “tangential” or “circumferential” refers to directions perpendicular to both the axial and radial directions. Therefore, as used herein, “radially inward” refers to a radial direction from the outer circumference of the gas turbine toward the centerline axis of the gas turbine, while “radially outward” refers to a radial direction from the centerline axis of the gas turbine toward the outer circumference of the gas turbine. As used herein, the terms “front,” “front portion,” and “forward” refer to a relatively upstream position in the airflow passing through or surrounding the component, while the terms “rear” and “rear portion” refer to a relatively downstream position in the airflow passing through or surrounding the component.

[0018] In some examples of open rotor engines, high vibrational loads are experienced during various phases of flight due to asymmetric propeller loading (e.g., P-factor or 1P loading). 1P loading (also known as + / -1P loading) refers to the movement or force on the blade caused by the blade's excitation frequency relative to the rotor's rotation. 1P loading typically occurs during high-power and high angle-of-attack operating conditions (e.g., takeoff). The 1P load experienced by the airfoil during engine operation can cause deflection of the airfoil (e.g., 1F deflection, etc.). This deflection generates loads and moments at the root of the affected blade and can lead to premature blade wear and failure. Some known turboprop or open rotor configurations make replacing individual blades difficult. In many examples, complex disassembly procedures must be performed to remove individual blades, increasing the time and cost required for gas turbine engine servicing.

[0019] Blade deflection can be reflected by a partial span shroud extending partially upwards from the blade face to the span portion of the blade. The partial span shroud may include fins and / or other components that interface with corresponding components of adjacent blades (e.g., adjacent, connected, etc.). The partial span shroud can mitigate the effects of IP loading and the resulting deflection. However, the prior art interface between partial span shrouds prevents blade pitch direction rotation. Therefore, prior art partial span shrouds prevent blade pitch control, thereby reducing the overall efficiency of the gas turbine engine.

[0020] The examples disclosed herein include blades with a partial span shield that enables the blade to rotate in the pitch direction. Examples disclosed herein include partial span shields that react to forces in the circumferential direction. In some examples disclosed herein, the partial span shield reacts to circumferential loads between adjacent circumferential directions but includes features that enable the pitch capability to be altered within the rotor airfoil. In some examples disclosed herein, the partial span shield includes a first section separated from a second section by a curved interface. In other examples disclosed herein, the partial span shield includes a tie rod connected to an adjacent airfoil via one or more slotted interfaces. In some examples disclosed herein, the partial span shield reduces the load borne at the root of the airfoil to which the partial span shield is connected. While the examples disclosed herein are primarily directed to open rotor engines, they are applicable to any suitable type of gas turbine (e.g., turbofan engines, turboprop engines, etc.). While the example partial span shields disclosed herein are typically positioned below 50% of the blade span (e.g., 25%, 30%, 40%, etc.), partial span shields can be positioned in any other suitable location.

[0021] The examples disclosed in this article can be applied to the design of both closed and open rotary engines. For illustrative purposes only. Figure 1 An example closed-loop rotary turbofan engine is shown, while Figure 2 An example open rotary engine is shown.

[0022] Figure 1 This is a cross-sectional view of a turbofan gas turbine engine from which the examples disclosed herein can be implemented. Referring now to the accompanying drawings, Figure 1 This is a schematic partial cross-sectional side view of an exemplary gas turbine engine 10, which can be combined with various examples of this disclosure. Engine 10 can be specifically configured as a gas turbine engine for an aircraft. Although further described herein as a turbofan engine, engine 10 can be defined as a turboshaft engine, a turboprop engine, or a turbojet gas turbine engine, including marine and industrial engines and auxiliary power units. Figure 1As shown, engine 10 has a longitudinal or axial centerline 12 extending through it for reference. An axial direction A extends in the same direction as the axial centerline 12 for reference. Engine 10 also defines an upstream end 99 and a downstream end 98 for reference. Typically, engine 10 may include a fan assembly 14 and a core engine 16 disposed downstream of the fan assembly 14. For reference, engine 10 defines an axial direction A, a radial direction R, and a circumferential direction C. Typically, axial direction A extends parallel to the axial centerline 12, radial direction R extends outward and inward from the axial centerline 12 in a direction orthogonal to axial direction A, and the circumferential direction extends 360° around the axial centerline 12.

[0023] The core engine 16 typically includes a generally tubular housing 18 defining an annular inlet 20. The housing 18 surrounds or at least partially forms, in a series flow relationship, a compressor section having a supercharger or low-pressure (LP) compressor 22, a high-pressure (HP) compressor 24, and a heat addition system 26; an expansion section or turbine section including a high-pressure (HP) turbine 28 and a low-pressure (LP) turbine 30; and an injection exhaust nozzle section 32. A high-pressure (HP) rotor shaft 34 drivesly connects the HP turbine 28 to the HP compressor 24. A low-pressure (LP) rotor shaft 36 drivesly connects the LP turbine 30 to the LP compressor 22. The LP rotor shaft 36 may also be connected to the fan shaft 38 of the fan assembly 14. In some examples, such as... Figure 1 As shown, for example in an indirect drive or gear drive configuration, the LP rotor shaft 36 can be connected to the fan shaft 38 via a reduction gear 40.

[0024] like Figure 1 As shown, the fan assembly 14 includes a plurality of fan blades 42 coupled to and extending radially outward from the fan shaft 38. An annular fan housing or nacelle 44 may circumferentially surround at least a portion of the fan assembly 14 and / or the core engine 16. Those skilled in the art will understand that the nacelle 44 may be configured to be supported relative to the core engine 16 by a plurality of circumferentially spaced outlet guide vanes or struts 46. Furthermore, at least a portion of the nacelle 44 may extend beyond the outer portion of the core engine 16 to define a fan flow passage 48 therebetween. However, it should be understood that various configurations of the engine 10 may omit the nacelle 44, or the nacelle 44 extending around the fan blades 42, for example, to provide… Figure 2 The open rotor or propeller fan structure of the engine 10 is depicted in the figure.

[0025] It should be understood that the combination of rotor shafts 34, 36, compressors 22, 24, and turbines 28, 30 defines the rotor assembly 90 of engine 10. For example, HP rotor shaft 34, HP compressor 24, and HP turbine 28 may define a high-speed or HP rotor assembly of engine 10. Similarly, the combination of LP rotor shaft 36, LP compressor 22, and LP turbine 30 may define a low-speed or LP rotor assembly of engine 10. Various examples of engine 10 may also include a fan shaft 38 and fan blades 42 as LP rotor assemblies. In some examples, engine 10 may further define a fan rotor assembly that is at least partially mechanically separated from the LP spool via fan shaft 38 and reduction gear 40. Further examples may define one or more intermediate rotor assemblies defined by an intermediate-pressure compressor, intermediate-pressure shaft, and intermediate-pressure turbine disposed between the LP rotor assembly and the HP rotor assembly (relative to a serial aerodynamic flow arrangement).

[0026] During operation of engine 10, an airflow, schematically indicated by arrow 74, enters inlet 76 of engine 10, defined by fan housing or nacelle 44. A portion of the air, schematically indicated by arrow 80, enters core engine 16 through an annular inlet 20 defined at least partially via outer casing 18. The airflow is provided in series via core flow path 70 through compressor, heat additive system, and expansion section. Airflow 80 is progressively compressed as it flows through successive stages of compressors 22, 24, as schematically indicated by arrow 82. Compressed air 82 enters heat additive system 26 and mixes with liquid and / or gaseous fuel, and is ignited to produce combustion gases 86. It should be understood that heat additive system 26 can be configured as any suitable system for generating combustion gases, including but not limited to detonation or knock combustion systems, or combinations thereof. Heat additive system 26 may include annular, canister, canister-annular, vortex, involute or vortex, rich, lean, rotary detonation, or pulse detonation configurations, or combinations thereof.

[0027] Combustion gas 86 releases energy to drive the rotation of the HP rotor assembly and LP rotor assembly before exiting from the injection exhaust nozzle section 32. The energy released from combustion gas 86 further drives the rotation of fan assembly 14, which includes fan blades 42. A portion of air 74 bypasses the core engine 16 and flows through fan flow passage 48, as schematically shown by arrow 78.

[0028] It should be understood that Figure 1 A dual-flow engine with a fan flow channel 48 and a core flow path 70 is depicted and described. Figure 1The example depicted has a nacelle 44 surrounding the fan blades 42 to provide noise reduction, blade shedding protection and / or other known benefits of the nacelle, and may be referred to herein as a “ducted fan,” or the entire engine 10 may be referred to as a “ducted engine.”

[0029] Figure 2 This is a schematic cross-sectional view of an example open rotor turbine engine according to one example of this disclosure. Specifically, Figure 2 The image shows an aviation three-flow turbofan engine, referred to in this article as "Three-flow Engine 100". Figure 2 The three-flow engine 100 can be mounted on an aircraft (such as a fixed-wing aircraft) and can generate thrust for propelling the aircraft. The architecture of the three-flow engine 100 provides three different airflows to generate thrust during operation. Figure 1 Unlike the engine 10 shown, the three-flow engine 100 includes a fan that is not guided by the nacelle or cowling, so it may be referred to herein as a “non-ducted fan”, or the entire engine 100 may be referred to as a “non-ducted engine”.

[0030] For reference, the three-flow engine 100 defines an axial direction A, a radial direction R, and a circumferential direction C. Furthermore, the engine 100 defines an axial centerline or longitudinal axis 112 extending along the axial direction A. Typically, the axial direction A extends parallel to the longitudinal axis 112, the radial direction R extends outward and inward from the longitudinal axis 112 in a direction perpendicular to the axial direction A, and the circumferential direction extends 360° around the longitudinal axis 112. The three-flow engine 100 extends, for example, along the axial direction A between a front end 114 and a rear end 116.

[0031] The three-flow engine 100 includes a core engine 120 and a fan section 150 positioned upstream therefrom. Typically, the core engine 120 includes a compressor section, a combustion section, a turbine section, and an exhaust section in a sequential flow order. Specifically, as... Figure 2As shown, the core engine 120 includes a core cowling 122 defining an annular core inlet 124. The core cowling 122 also surrounds a low-pressure system and a high-pressure system. In some examples, the core cowling 122 may surround and support a turbocharger or low-pressure (“LP”) compressor 126 for pressurizing air entering the core engine 120 through the core inlet 124. A high-pressure (“HP”) multistage axial compressor 128 receives pressurized air from the LP compressor 126 and further increases the air pressure. The pressurized air flows downstream to a combustor 130, where fuel is injected into the pressurized air flow and ignited to increase the temperature and energy level of the pressurized air. It should be understood that, as used herein, the terms “high / low speed” and “high / low pressure” are used interchangeably with respect to a high-pressure / high-speed system and a low-pressure / low-speed system. Furthermore, it should be understood that the use of the terms “high” and “low” in the same context to distinguish between two systems does not imply any absolute speed and / or pressure values.

[0032] High-energy combustion products flow downstream from combustor 130 to high-pressure turbine 132. High-pressure turbine 132 drives high-pressure compressor 128 via high-pressure shaft 136. In this respect, high-pressure turbine 132 is drivably coupled to high-pressure compressor 128. High-energy combustion products then flow to low-pressure turbine 134. Low-pressure turbine 134 drives components of low-pressure compressor 126 and fan section 150 via low-pressure shaft 138. In this respect, low-pressure turbine 134 is drivably coupled to components of low-pressure compressor 126 and fan section 150. In this example, LP shaft 138 is coaxial with HP shaft 136. After driving each of turbines 132, 134, combustion products exit core engine 120 through core exhaust nozzle 140 to generate propulsive thrust. Therefore, core engine 120 defines a core flow path or core conduit 142 extending between core inlet 124 and core exhaust nozzle 140. The core duct 142 is an annular duct that is typically located in the radial direction R inside the core fairing 122.

[0033] Fan section 150 includes fan 152, which in this example is the primary fan. For Figure 2 In the example shown, fan 152 is an open rotor or non-ducted fan. However, in other examples, fan 152 may be guided, for example, by a fan housing or nacelle circumferentially surrounding fan 152. As shown, fan 152 includes an array of fan blades 154. Figure 2 (Only one is shown). The fan blade 154 is rotatable, for example, about the longitudinal axis 112. As described above, the fan 152 is drivenly connected to the low-pressure turbine 134 via the LP shaft 138. For example, in a direct drive configuration, the fan 152 can be directly connected to the LP shaft 138. Alternatively, as... Figure 2As shown, for example in an indirect drive or gear drive configuration, the fan 152 can be connected to the LP shaft 138 via a reduction gearbox 155.

[0034] Furthermore, the fan blades 154 can be arranged at equal intervals around the longitudinal axis 112. Each blade 154 has a root and a tip, and a span defined between them. Each blade 154 defines a central blade axis 156. In this example, each blade 154 of the fan 152 is capable of rotating about its respective central blade axis 156, for example, in unison with each other. One or more actuators 158 can be controlled to pitch the blades 154 about their respective central blade axes 156. However, in other examples, each blade 154 may be fixed or unable to pitch about its central blade axis 156.

[0035] Fan section 150 also includes a fan guide vane array 160, which includes fan guide vanes 162 arranged around a longitudinal axis 112. Figure 2 (Only one is shown in the image). In this example, the fan guide vane 162 is not capable of rotating about the longitudinal axis 112. Each fan guide vane 162 has a root and a tip, and a span defined between them. The fan guide vane 162 can be as follows: Figure 2 The fan guide vanes 162 shown are not obscured, or may be obscured, for example, by an annular shroud spaced outward from the tips of the fan guide vanes 162 in the radial direction R or attached to the fan guide vanes 162. Each fan guide vane 162 defines a central blade axis 164. In this example, each fan guide vane 162 of the fan guide vane array 160 is capable of rotating about its respective central blade axis 164, for example, in unison with each other. One or more actuators 166 may be controlled to pitch the fan guide vanes 162 about their respective central blade axes 164. However, in other examples, each fan guide vane 162 may be fixed or unable to pitch about its central blade axis 164. The fan guide vanes 162 are mounted to the fan shroud 170.

[0036] like Figure 2As shown, in addition to the non-ducted fan 152, a ducted fan 184 is also included behind fan 152, such that the three-flow engine 100 includes both ducted and non-ducted fans, both used to generate thrust by the movement of air without passing through the core engine 120. The ducted fan 184 is shown positioned at approximately the same axial location as the fan guide vane 162 and radially inward of the fan guide vane 162. Alternatively, the ducted fan 184 may be located between the fan guide vane 162 and the core duct 142, or further forward of the fan guide vane 162. The ducted fan 184 may be driven by a low-pressure turbine 134 (e.g., coupled to the LP shaft 138) or by any other suitable rotational source, and may be used as the first stage of a supercharger or may operate independently.

[0037] A fan fairing 170 annularly surrounds at least a portion of the core fairing 122 and is generally positioned radially R outside the core fairing 122. Specifically, a downstream section of the fan fairing 170 extends above the front portion of the core fairing 122 to define a fan flow path or fan duct 172. Incoming air may enter the fan duct 172 through a fan duct inlet 176 and may exit through a fan exhaust nozzle 178 to generate propulsive thrust. The fan duct 172 is an annular duct generally positioned radially R outside the core duct 142. Stationary struts 174 may each have an aerodynamic profile to guide airflow therefrom. Other struts besides the stationary struts 174 may be used to connect and support the fan fairing 170 and / or the core fairing 122. In many examples, the fan duct 172 and the core fairing 122 may at least partially coexist (generally axially) on opposite sides (e.g., opposite radial sides) of the core fairing 122. For example, the fan duct 172 and the core fairing 122 may each extend directly from the leading edge 144 of the core fairing 122, and may extend together in a generally axial manner on the opposite radial sides of the core fairing.

[0038] The three-flow engine 100 also defines or includes an inlet duct 180. The inlet duct 180 extends between the engine inlet 182 and the core inlet 124 / fan duct inlet 176. The engine inlet 182 is generally defined at the front end of the fan cowl 170 and is positioned along the axial direction A between the fan 152 and the fan guide vane array 160. The inlet duct 180 is an annular duct positioned radially R inside the fan cowl 170. Air flowing downstream along the inlet duct 180 is diverted by a splitter or leading edge 144 of the core cowl 122 into the core duct 142 and the fan duct 172, and is not necessarily uniform. The inlet duct 180 is wider in the radial direction R than the core duct 142. The inlet duct 180 is also wider in the radial direction R than the fan duct 172.

[0039] Figure 3 It includes the example section with a span of 300. Figure 1 Front view of fan 152. Figure 3 In the fan 152, there is a disk 301 having a first blade 302 and a second blade 304, and a portion 306 of a partially spanned shroud 300. Figure 3 In the middle, part 306 includes example interface 308.

[0040] exist Figure 3 In this configuration, the first blade 302 and the second blade 304 are coupled to the disk 301 (e.g., via dovetail grooves and corresponding grooves of the disk 301). In other examples, the disk 301 and the blades 302, 304 may be integral (e.g., cast, manufactured via negative manufacturing, manufactured via additive manufacturing, etc.). Although the disk 301 is described as a disk of an open rotor, in other examples, the teachings of this disclosure can be applied to any other suitable type of rotor (e.g., Figure 1 Fan assembly 14, etc.

[0041] The interface 308 of portion 306 of the partial span shield 300 serves as a joint to react to circumferential loads between blades 302 and 304. Specifically, interface 308 allows circumferential loads to react between the blades to reduce deflection caused by P-loads during operation. In some examples, interface 308 does not react to moments (e.g., pitch moments) between blades 302 and 304. Thus, the partial span shield 300 allows the blades to rotate about their radial axes, which aids in pitch control of blades 302 and 304.

[0042] exist Figure 3 In this configuration, a partial span shield 300 is positioned at 30% of the span of blades 302 and 304. In some examples, the position of the partial span shield 300 reduces the deflection of blades 302 and 304 near disk 301 to mitigate the effect of P-load on blades 302 and 304. In other examples, the partial span shield 300 may be positioned at any other suitable location on blades 302 and 304 (e.g., less than 50% span, 25% span, etc.).

[0043] Figure 4 yes Figure 3 A top view of part 306 of the span shield 300. (See attached image.) Figure 4 As shown in the example, interface 308 includes a first portion 400 having a first surface 402 and a second portion 404 having a second surface 406. Figure 4In this configuration, a first portion 400 is connected to a first blade 302, and a second portion 404 is connected to a second blade 304. For example, portions 400 and 404 may be connected to the respective blades of blades 302 and 304 via one or more welding processes. In other examples, the first portion 400 and the first blade 302 are integral and / or the first portion 400 and the first blade 302 are integral. In some such examples, blades 302 and 304 and the corresponding portions 400 and 404 may be manufactured via additive manufacturing, negative manufacturing, casting, etc.

[0044] exist Figure 4 In the example shown, surfaces 402 and 406 form interface 308. In Figure 4 In this example, interface 308 is a circular interface. In other examples, interface 308 can have any other suitable shape. Figure 4 In this example, the first surface 402 is a convex surface and the second surface 406 is a concave surface. In other examples, the first surface 402 is a concave surface and the second surface 406 is a convex surface. Figure 4 In this configuration, portions 400 and 404 are configured to be adjacent when the fan 152 is in motion (e.g., in a gas turbine engine 10, gas turbine engine 100, etc.). Therefore, when the engine is cooled (e.g., not in operation), surfaces 402 and 406 are not adjacent. When the fan is in operation, blades 302 and 304 and portions 400 and 404 expand due to mechanical and thermal loads, causing surfaces 402 and 406 to come into contact with each other. In such examples, the abutment of the first portion 400 and the second portion 404 allows circumferential loads to react between the first portion 400 and the second portion 404, and between the first blade 302 and the second blade 304.

[0045] The interface 308 formed by surfaces 402 and 406 allows portions 400 and 404 to rotate relative to the other portions, which allows rotation of the blades 302 and 304 in the relative pitch direction. In some examples, surfaces 402 and 406 may be coated with anti-friction bushings (e.g., PTFE, plastic, polymer, etc.) and / or otherwise manufactured (e.g., polished, surface-treated, etc.) to reduce friction associated with the abutment and relative movement of surfaces 402 and 406. In the example shown above, the curved interface is an arc. In other examples, any other suitable type of shape may be used (e.g., elliptical, hyperbolic, parabolic, cubic, etc.). Thus, the partial span shield 300 facilitates pitch control using fan 152 while preventing undesirable deflection of blades 302 and 304 due to 1P loading.

[0046] Figure 5 This is a top view of the replacement portion of the 500-span shield. Figure 5 In the middle, the replacement part of the span shield 500 is in Figure 3 It extends between the first blade 302 and the second blade 304. The partial span shroud 500 includes a first pin 504 and a second pin 506, which are respectively disposed in a first groove 508 and a second groove 510. Figure 5 In the middle, grooves 508 and 510 are formed in the first part 512 and the second part 514.

[0047] exist Figure 5 In this configuration, the pull rod 502 is rotatably connected to the first blade 302 via a first pin 504 disposed in a first groove 508, and is rotatably connected to the second blade 304 via a second pin 506 disposed in a second groove 510. Figure 5 In this configuration, pins 504 and 506 are disposed within slots 508 and 510 along the radial axis R. During operation, pins 504 and 506 can circumferentially translate within slots 508 and 510, respectively, thereby promoting rotation of blades 302 and 304 in the pitch direction (e.g., radial direction). Tie rod 502 reacts to the circumferential load between blades 302 and 304. Figure 5 In this example, the tie rod 502 is cylindrical and has a solid cross-section. In other examples, the tie rod 502 may have any other suitable shape and / or cross-section (e.g., hollow, etc.).

[0048] exist Figure 5 In this configuration, the first portion 512 is connected to the first blade 302 and the second portion 514 is connected to the second blade 304. For example, portions 512 and 514 may be connected to the respective blades of blades 302 and 304 via one or more welds or the like. In other examples, the first portion 512 and the first blade 302 are integral and / or the second portion 514 and the first blade 302 are integral. In some such examples, blades 302 and 304 and the corresponding portions 512 and 514 may be manufactured via additive manufacturing, negative manufacturing, casting, or the like.

[0049] "Comprising" and "including" (and all their forms and tenses) are used herein as open-ended terms. Therefore, whenever a claim uses any form of "comprising" or "including" (e.g., including, comprising, having, etc.) in the preamble or in any type of claim statement, it should be understood that additional elements, terms, etc., may be present without exceeding the scope of the corresponding claim or statement. As used herein, when the phrase "at least" is used as a transitional term, for example, in the preamble of a claim, it is open-ended in the same way that the terms "comprising" and "including" are open-ended. The term "and / or," when used, for example, in the form of A, B, and / or C, refers to any combination or subset of A, B, and C, such as (1) A alone, (2) B alone, (3) C alone, (4) A and B, (5) A and C, (6) B and C, or (7) A and B and C. As used herein in the context of describing structures, components, items, objects, and / or things, the phrase "at least one of A and B" is intended to refer to any implementation including (1) at least one A, (2) at least one B, or (3) at least one A and at least one B. Similarly, as used herein in the context of describing structures, components, items, objects, and / or things, the phrase "at least one of A or B" is intended to refer to any implementation including (1) at least one A, (2) at least one B, or (3) at least one A and at least one B. As used herein in the context of describing the conduct or execution of processes, instructions, actions, activities, and / or steps, the phrase "at least one of A and B" is intended to refer to any implementation including (1) at least one A, (2) at least one B, or (3) at least one A and at least one B. Similarly, as used herein in the context of describing the conduct or execution of processes, instructions, actions, activities, and / or steps, the phrase "at least one of A or B" is intended to refer to any implementation including (1) at least one A, (2) at least one B, or (3) at least one A and at least one B.

[0050] As used herein, singular references (e.g., “a,” “an,” “first,” “second,” etc.) do not exclude plurals. As used herein, the term “a” or “an” refers to one or more of those objects. The terms “a” (or “an”), “one or more,” and “at least one” are used interchangeably herein. Furthermore, although listed separately, multiple means, elements, or method actions can be implemented by, for example, the same entity or object. Moreover, while individual features may be included in different examples or claims, these may be combined, and inclusion in different examples or claims does not imply that the combination of features is impractical and / or disadvantageous.

[0051] Further aspects of the invention are provided by the subject matter of the following clauses:

[0052] Example 1 includes a device comprising: a first portion extending from a first airfoil, the first portion including a first surface that is convex; a second portion extending from a second airfoil toward the first airfoil, the second portion including a second surface, the first airfoil being circumferentially adjacent to the second airfoil, the second surface being concave; and an interface formed by the first surface and the second surface, the interface reacting to a circumferential load between the first airfoil and the second airfoil.

[0053] Example 2 includes a device according to any of the foregoing clauses, wherein the interface does not react to the torque between the first airfoil and the second airfoil.

[0054] Example 3 includes a device according to any of the preceding clauses, wherein the interface is a circular connector.

[0055] Example 4 includes a device according to any of the preceding clauses, wherein at least one of the first surface or the second surface includes a friction-reducing bushing.

[0056] Example 5 includes the device described in any of the preceding clauses, wherein the first airfoil and the first portion are integral.

[0057] Example 6 includes the device according to any of the foregoing clauses, wherein at least one of the first portion or the second portion is disposed at a span less than 35% of the first airfoil.

[0058] Example 7 includes a device according to any of the preceding clauses, wherein the first airfoil and the second airfoil are coupled to an open rotor.

[0059] Example 8 includes a gas turbine engine comprising: a disk; a first airfoil coupled to the disk; a second airfoil coupled to the disk and circumferentially adjacent to the first airfoil; and a partial span shield comprising: a first portion extending from the first airfoil and including a first surface that is convex; a second portion extending from the second airfoil toward the first airfoil and including a second surface that is concave, the first airfoil being circumferentially adjacent to the second airfoil; and an interface formed by the first surface and the second surface, the interface reacting to a circumferential load between the first airfoil and the second airfoil.

[0060] Example 9 includes a gas turbine engine according to any of the preceding clauses, wherein the interface does not react to the torque between the first airfoil and the second airfoil.

[0061] Example 10 includes a gas turbine engine according to any of the foregoing clauses, wherein the interface is a circular connector.

[0062] Example 11 includes a gas turbine engine according to any of the preceding clauses, wherein at least one of the first surface or the second surface includes a friction-reducing bushing.

[0063] Example 12 includes a gas turbine engine according to any of the foregoing clauses, wherein the first airfoil and the first portion are integral.

[0064] Example thirteen includes a gas turbine engine according to any of the preceding clauses, wherein at least one of the first portion or the second portion is disposed at a span less than 35% of the first airfoil.

[0065] Example fourteen includes a gas turbine engine according to any of the preceding clauses, wherein the disk is an open rotor disk.

[0066] Example 15 includes a device comprising: a first portion extending from a first airfoil; a second portion extending from a second airfoil toward the first airfoil; and a structural member comprising: a first end rotatably connected to the first portion; and a second end rotatably connected to the second portion.

[0067] Example 16 includes a device according to any of the preceding clauses, wherein: the structural member reacts to a circumferential load between the first airfoil and the second airfoil; and the structural member does not react to a moment between the first airfoil and the second airfoil.

[0068] Example 17 includes the device according to any of the preceding clauses, further comprising: a first groove formed in the first portion; a second groove formed in the second portion; a first pin disposed in the first groove, the first pin rotatably connecting the first portion to the structural member; and a second pin disposed in the second groove, the second pin rotatably connecting the second portion to the structural member.

[0069] Example 18 includes a device according to any of the preceding clauses, wherein the first airfoil and the second airfoil extend radially from the disk, and the first pin and the second pin extend along a radial axis.

[0070] Example 19 includes the device according to any of the preceding clauses, wherein at least one of the first portion or the second portion is disposed at a span less than 35% of the first airfoil.

[0071] Example 20 includes the device according to any of the preceding clauses, wherein the first airfoil and the first portion are integral.

[0072] The following claims are incorporated herein by reference, and each claim exists independently as a separate embodiment of this disclosure.

Claims

1. A device, characterized in that, include: The first part extends from the first airfoil and includes a first surface that is convex. The second part extends from the second airfoil toward the first airfoil, and the second part includes a second surface that is circumferentially adjacent to the first airfoil and is concave. as well as An interface, which is an arcuate interface formed by the first surface and the second surface, reacts to the circumferential load between the first airfoil and the second airfoil, and allows the first airfoil and the second airfoil to rotate in the relative pitch direction while preventing the first airfoil and the second airfoil from deflecting undesirably when subjected to load.

2. The device according to claim 1, characterized in that, in, The interface does not react to the torque between the first airfoil and the second airfoil.

3. The device according to claim 2, characterized in that, in, The interface is a circular connector.

4. The device according to claim 1, characterized in that, in, At least one of the first surface or the second surface includes a friction-reducing bushing.

5. The device according to claim 1, characterized in that, in, The first airfoil and the first part are integral.

6. The device according to claim 1, characterized in that, in, At least one of the first portion or the second portion is positioned at a distance less than 35% of the span of the first airfoil.

7. The device according to claim 1, characterized in that, in, The first airfoil and the second airfoil are connected to the open rotor.

8. A gas turbine engine, characterized in that, include: plate; A first airfoil is connected to the disk; A second airfoil is connected to the disk and is circumferentially adjacent to the first airfoil; as well as Partial span protective cover, the partial span protective cover comprising: The first portion extends from the first airfoil and includes a first surface that is convex. The second portion extends from the second airfoil toward the first airfoil, and includes a second surface circumferentially adjacent to the first airfoil, the second surface being concave; and An interface, which is an arcuate interface formed by the first surface and the second surface, reacts to the circumferential load between the first airfoil and the second airfoil, and allows the first airfoil and the second airfoil to rotate in the relative pitch direction while preventing the first airfoil and the second airfoil from deflecting undesirably when subjected to load.

9. The gas turbine engine according to claim 8, characterized in that, in, The interface does not react to the torque between the first airfoil and the second airfoil.

10. The gas turbine engine according to claim 9, characterized in that, in, The interface is a circular connector.

11. The gas turbine engine according to claim 8, characterized in that, in, At least one of the first surface and the second surface includes a friction-reducing bushing.

12. The gas turbine engine according to claim 8, characterized in that, in, The first airfoil and the first part are integral.

13. The gas turbine engine according to claim 8, characterized in that, in, At least one of the first portion or the second portion is positioned at a distance less than 35% of the span of the first airfoil.

14. The gas turbine engine according to claim 8, characterized in that, in, The disk is an open rotor disk.

15. A device, characterized in that, include: A first portion, the first portion extending from a first airfoil, the first portion including a first groove; The second part extends from the second airfoil toward the first airfoil; as well as Structural components, the structural components including: The first end is rotatably connected to the first part; The second end, which is rotatably connected to the second part; and A first pin is disposed in the first groove, the first pin rotatably connects the first portion to the structural member, and the first pin is capable of circumferential translation within the first groove.

16. The device according to claim 15, characterized in that, in: The structural member reacts to the circumferential load between the first airfoil and the second airfoil; and The structural component does not react to the torque between the first airfoil and the second airfoil.

17. The device according to claim 15, characterized in that, in, The second portion defines a second groove and further includes a second pin disposed in the second groove, the second pin rotatably connecting the second portion to the structural member.

18. The device according to claim 17, characterized in that, in, The first airfoil and the second airfoil extend radially from the disk, and the first pin and the second pin extend along the radial axis.

19. The device according to claim 15, characterized in that, in, At least one of the first portion or the second portion is positioned at a distance less than 35% of the span of the first airfoil.

20. The device according to claim 15, characterized in that, in, The first airfoil and the first part are integral.

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