Assembly for controlling the clearance between a rotor blade and a composite casing in a turbine engine

Abstract

An assembly to control a clearance between a rotor blade and a casing in a turbine engine includes a casing having an annular casing body and a liner positioned adjacent an inner radial surface of the annular casing body, the casing being spaced apart from the rotor blade in a radial direction by the clearance. A coefficient of thermal expansion (CTE) of the annular casing body is lower than a CTE of the liner and a CTE of the rotor blade. The liner is provided to control the clearance between the rotor blade and the casing.

Description

[0001] Cross-reference to related applications

[0002] This application claims the benefit of Polish patent application No. P.451169, filed on February 10, 2025, which is incorporated herein by reference in its entirety. Technical Field

[0003] This disclosure generally relates to components for controlling the gap between rotor blades and composite housing in a turbine engine. Background Technology

[0004] Turbine engines used in aircraft typically consist of rotor blades housed in a casing. The rotor blades are usually made of metals such as titanium, while the casing is made of composite materials. Using a composite casing instead of a metal casing allows for a reduction in system weight. The efficiency of a turbine engine depends at least in part on the clearance or gap between the casing and the rotor blades. If the clearance is too large, an excessive portion of the engine airflow will pass through the gap instead of interacting with the rotating blades, resulting in reduced engine efficiency. If the clearance is too small, interference may occur between the rotor blades and the casing, leading to damage to one or more surfaces of the rotor blades or casing.

[0005] During turbine engine operation, rotor blades tend to thermally expand or contract, while the casing, due to its composite material structure, does not thermally expand or contract, or contracts only to a minimal extent, in the presence of a thermal gradient. Due to the thermal expansion of the rotor blades, a gap is provided between the blade tips and the casing to accommodate this expansion. Higher performance can be achieved by providing a smaller or tighter gap between the blade tips and the casing. Attached Figure Description

[0006] Features and advantages will become apparent from the following more specific description of various exemplary embodiments, as illustrated in the accompanying drawings, wherein similar reference numerals generally denote the same elements, functionally similar elements, or structurally similar elements, or both.

[0007] Figure 1 This is a schematic cross-sectional view of a turbine engine taken along the longitudinal centerline axis of the turbine engine according to an embodiment of the present disclosure.

[0008] Figure 2 According to embodiments of this disclosure Figure 1 A cross-sectional view of a portion of the inlet duct of the turbine engine shown. Figure 1 Detail 2 in the text.

[0009] Figure 3This is a cross-sectional view of an assembly including a fan housing, taken along the radial direction of a turbine engine according to an embodiment of the present disclosure. The fan housing has an annular housing body, a metal bushing, and a plurality of metal spring arms.

[0010] Figure 4A It is according to an embodiment of the present disclosure along the axial direction of the turbine engine. Figure 3 An axial view of the metal bushing taken at line 4A-4A.

[0011] Figure 4B This is an axial view of a metal bushing along the axial direction of a turbine engine, according to another embodiment of this disclosure.

[0012] Figure 4C This is an axial view of a metal bushing along the axial direction of a turbine engine, according to yet another embodiment of this disclosure.

[0013] Figure 5 According to embodiments of this disclosure Figure 3 A cross-sectional view of the metal bushing taken at line 5-5.

[0014] Figure 6 This is a cross-sectional view of an assembly including a fan housing, taken along the radial direction of a turbine engine according to an embodiment of the present disclosure. The fan housing has an annular housing body, a metal bushing, and a plurality of spacers.

[0015] Figure 7 This is a cross-sectional view of a metal bushing according to an embodiment of the present disclosure. Detailed Implementation

[0016] The features, advantages, and embodiments of this disclosure will be set forth or apparent from consideration of the following detailed description, accompanying drawings, and claims. Furthermore, the foregoing summary and the following detailed description of this disclosure are exemplary and intended to provide further explanation, without limiting the scope of the claimed disclosure.

[0017] Various embodiments of this disclosure are discussed in detail below. While specific embodiments are discussed, they are for illustrative purposes only. Those skilled in the art will recognize that other components and constructions can be used without departing from this disclosure.

[0018] As used herein, the terms “first” and “second” are used interchangeably to distinguish one component from another and are not intended to indicate the location or importance of the individual components.

[0019] 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.

[0020] Unless otherwise stated, when used with compressor, turbine, shaft, or spool components, the terms “low” and “high,” or their respective comparatives (e.g., “lower” and “higher,” where applicable), refer to relative pressure or relative speed within the engine, or both. For example, a “low-speed” component defines a component configured to operate at a speed (such as the maximum permissible speed) lower than that of a “high-speed” component of the engine. Alternatively, unless otherwise stated, the foregoing terms may be understood in their superlative sense. For example, a “low-pressure turbine” may refer to the lowest maximum pressure within the turbine section, while a “high-pressure turbine” may refer to the highest maximum pressure within the turbine section. The terms “low” or “high” in such aspects may additionally or alternatively be understood relative to the minimum permissible speed or pressure, or relative to the minimum or maximum permissible speed or pressure of normal, desired, steady-state, etc., operation of the engine, or any combination thereof.

[0021] Unless otherwise specified herein, the terms “connection,” “fixation,” “attachment,” “linkage,” etc., refer to both direct connection, fixation, attachment, or linking, and indirect connection, fixation, attachment, or linking through one or more intermediate components or features.

[0022] Unless the context clearly indicates otherwise, the singular forms “a,” “a,” and “the” include plural references.

[0023] As used herein, the term "axial" refers to the direction and orientation extending substantially parallel to the longitudinal centerline of the turbine engine. Furthermore, the terms "radial" and "radially" refer to the direction and orientation extending substantially perpendicular to the longitudinal centerline of the turbine engine. Additionally, as used herein, the term "circumferential" refers to the direction and orientation extending in an arc around the longitudinal centerline of the turbine engine.

[0024] As used herein, the term "composite material" refers to component materials having two or more constituent materials. A composite material can be a combination of at least two or more metallic or nonmetallic elements or materials. Examples of composite materials include, but are not limited to, polymer matrix composites (PMC), ceramic matrix composites (CMC), metal matrix composites (MMC), carbon fibers, polymeric resins, thermoplastic resins, bismaleimide (BMI) materials, polyimide materials, epoxy resins, glass fibers, and silicon matrix materials. Composite materials can be formed from matrix materials and reinforcing elements, such as fibers (referred to herein as reinforcing fibers).

[0025] As used herein, "reinforcing fiber" may include, for example, glass fiber, carbon fiber, steel fiber, or para-aramid fiber, such as those available from DuPont of Wilmington, Delaware. The reinforcing fibers can be in the form of fiber bundles, which include multiple fibers forming the bundles. The polymer matrix material can include, for example, thermosetting resins, bismaleimide (BMI) materials, polyimide materials, or thermoplastic resins.

[0026] Composite materials can be used to form composite components (e.g., composite airfoils). As used herein, a “composite component” refers to a structure or component that comprises any suitable composite material. Composite components (such as composite airfoils (e.g., composite fan blades)) may comprise several layers or several plies of composite material (composite plies). The stiffness, material, and dimensions of the layers or plies can be varied to achieve a desired composite component or composite portion of a component having a predetermined weight, size, stiffness, and strength. One or more adhesives may be used to form or bond the composite component. Adhesives may include resins and phenolic resins, where the adhesive may require curing at elevated temperatures or other hardening techniques.

[0027] As may be used herein, PMC refers to a class of materials. PMC materials can be prepregs. A prepreg is a reinforcing material (e.g., reinforcing fibers) pre-impregnated with a polymer matrix material (e.g., a thermoplastic resin). Non-limiting examples of processes for producing thermoplastic prepregs include: hot melt prepreg, in which the fiber reinforcing material is drawn through a molten bath of resin; and powder prepreg, in which the resin is deposited onto the fiber reinforcing material, as a non-limiting example, electrostatically deposited onto the fiber reinforcing material, and then adhered to the fibers, as a non-limiting example, in an oven or with the aid of heated rollers.

[0028] Resins used in PMC matrix materials are generally classified as thermosetting or thermoplastic resin polymers. Thermoplastic resin polymers are typically categorized as polymers that can repeatedly soften and flow upon heating and harden upon sufficient cooling due to physical rather than chemical changes. Well-known examples of thermoplastic resin polymers include nylon, thermoplastic polyesters, polyaryletherketones (PAEs), and polycarbonate resins. Specific examples of high-performance thermoplastic resins envisioned for aerospace applications include polyetheretherketone (PEEK), polyetherketoneketone (PEKK), polyetherimide (PEI), polyaryletherketone (PAEK), and polyphenylene sulfide (PPS). Conversely, thermosetting resins do not undergo significant softening upon heating once fully cured into a rigid solid; instead, they thermally decompose upon sufficient heating. Well-known examples of thermosetting resin polymers include epoxy resins, bismaleimide (BMI), and polyimide resins.

[0029] Instead of using prepregs with thermoplastic polymers, another non-limiting example utilizes woven fabrics. Woven fabrics may include, but are not limited to, dry carbon fibers woven together with thermoplastic polymer fibers or filaments. Non-prepreg woven structures can be fabricated in a similar manner. With this method, the fiber volume of the part can be customized by specifying the relative concentrations of the woven or braided thermoplastic fibers and reinforcing fibers. Furthermore, different types of reinforcing fibers can be woven or braided together at different concentrations to customize the properties of the part. For example, glass fibers, carbon fibers, and thermoplastic fibers can all be woven together at different concentrations to customize the properties of the part. Carbon fibers provide the strength of the system, can be incorporated into glass fibers to enhance impact characteristics—a design feature of parts located near the engine inlet—and thermoplastic fibers provide bonding for the reinforcing fibers.

[0030] In yet another non-limiting example, resin transfer molding (RTM) can be used to form at least a portion of a composite part. Typically, RTM involves applying dry fibers to a mold or cavity. The dry fibers can include prepreg, braided material, woven material, or any combination thereof. Resin can be pumped into or otherwise supplied to the mold or cavity to impregnate the dry fibers. The impregnated fibers combined with the resin are then cured and removed from the mold. The composite part may require post-curing treatment upon removal from the mold. RTM can be a vacuum-assisted process. That is, air in the cavity or mold can be removed and replaced with resin before heating or curing. The placement of the dry fibers can be manual or automatic. The dry fibers can be shaped to form the composite part or guide the resin. Optionally, additional layers or reinforcing layers of materials different from the dry fibers can be included or added before heating or curing.

[0031] As used herein, CMC refers to a class of materials having reinforcing fibers within a ceramic matrix. Typically, the reinforcing fibers provide structural integrity to the ceramic matrix. Some examples of reinforcing fibers may include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), non-oxide carbon-based materials (e.g., carbon), oxide ceramics (e.g., silicon oxycarbide, silicon oxynitride, alumina (Al₂O₃), silicon dioxide (SiO₂), aluminosilicates (such as mullite), or mixtures thereof), or mixtures thereof.

[0032] Examples of ceramic matrix materials may include, but are not limited to, non-oxide silicon-based materials (e.g., silicon carbide, silicon nitride, or mixtures thereof), oxide ceramics (e.g., silicon oxycarbide, silicon oxynitride, alumina (Al₂O₃), silicon dioxide (SiO₂), aluminosilicates, or mixtures thereof), or mixtures thereof. Optionally, ceramic particles (e.g., oxides of Si, Al, Zr, Y, and combinations thereof) and inorganic fillers (e.g., pyrophyllite, wollastonite, mica, talc, kyanite, and montmorillonite) may also be included within the ceramic matrix.

[0033] Typically, a particular CMC can be referred to as a combination of its fiber type / matrix type. For example, C / SiC is carbon fiber reinforced silicon carbide, SiC / SiC is silicon carbide fiber reinforced silicon carbide, SiC / SiN is silicon carbide fiber reinforced silicon nitride, and SiC / SiC-SiN is a silicon carbide fiber-reinforced silicon carbide / silicon nitride matrix mixture, etc. In other examples, a CMC may consist of a matrix comprising an oxide-based material (such as alumina (Al₂O₃), silicon dioxide (SiO₂), aluminosilicates, and mixtures thereof) and reinforcing fibers. Aluminosilicates may include crystalline materials (e.g., mullite (3Al₂O₃·2SiO₂)) as well as glassy aluminosilicates.

[0034] In some non-limiting examples, the reinforcing fibers may be bundled (e.g., forming fiber bundles) or coated, or both, before being incorporated into the matrix. The fiber bundles may be impregnated with a slurry composition before or after the formation of the preform. The preform may then undergo heat treatment and subsequent chemical treatment to obtain a component formed from a CMC material having a desired chemical composition. For example, the preform may undergo curing or burn-out to produce a high coke residue in the preform and subsequently melt infiltration with silicon, or undergo curing or pyrolysis to produce a silicon carbide matrix in the preform and subsequently chemical vapor infiltration with silicon carbide. Additional steps may be taken to enhance the densification of the preform, either before or after chemical vapor infiltration, by injecting the preform with a liquid resin or polymer followed by a heat treatment step to fill the voids with silicon carbide. The CMC material used herein may be formed using any known or later developed method (including, but not limited to, melt infiltration, chemical vapor infiltration, polymer impregnation pyrolysis (PIP), or any combination thereof).

[0035] As used herein, the term "metal" refers to materials that include metals (such as, but not limited to, titanium, iron, aluminum, stainless steel, and nickel alloys). Metallic materials or alloys can be combinations of at least two or more elements or materials (at least one of which is a metal).

[0036] As used herein, an alloy is "based on" a particular element when that element constitutes the largest weight percentage (by total weight) of all the elements contained in the alloy. For example, an iron-based alloy has a higher weight percentage of iron than any other single element present in the alloy.

[0037] As mentioned above, certain components of gas turbine engines, particularly those used in aircraft, can be made of composite materials. These components can include various housings and shell structures. For example, a rotating airfoil assembly (such as a fan or compressor rotor) includes rotating blades and can have a housing surrounding the blades. Including a composite housing provides significant weight reduction for use in aircraft turbine engines. Composite materials include reinforcing fiber bundles, such as reinforcing fiber bundles formed from carbon fibers. Carbon fibers have a very low coefficient of thermal expansion (CTE), almost zero in many engineering applications; therefore, composite housings made of carbon fiber composites typically expand very little during engine operation. When the housing is used for rotor blades where the blades are also composite materials, the blades and housing expand similarly during changes in operating conditions. However, when the housing is used for rotor blades with metallic materials, a larger radial clearance is maintained between the rotor blades and the composite housing to prevent the rotor blades from rubbing against the housing during engine operation, as the rotor blades (metallic rotor blades) expand more than the housing. This larger radial clearance will reduce engine performance and fuel efficiency. The efficiency of a turbine engine depends at least in part on the clearance or void between the housing and the rotor blades. If the gap is too large, too much of the engine flow will pass through the gap instead of interacting with the rotating blades, resulting in reduced engine efficiency. If the gap is too small, interference may occur between the rotor blades and the housing, leading to damage to one or more surfaces of the rotor blades or housing.

[0038] This disclosure provides an assembly for controlling the clearance between rotor blades and a housing, wherein the housing has a lower CTE than the rotor blades. The rotor blades are typically made of metal (such as titanium), while the housing is typically made of a composite material. Using a composite housing instead of a metal housing allows for a reduction in system weight. Rotor blades tend to thermally expand in the presence of a thermal gradient. However, the housing does not thermally expand, or only slightly expands, in the presence of a thermal gradient. Due to the thermal expansion of the rotor blades, the clearance between the blade tips and the housing is controlled to accommodate this thermal expansion.

[0039] A better match is desired for the CTE system, allowing the housing and rotor blades to expand at similar rates, thereby maintaining a tighter or smaller clearance. In this embodiment, a metal bushing is provided for the composite housing to address thermal expansion. The use of the metal bushing provides the ability to substantially match the thermal expansion of the rotor blades. Therefore, a tighter clearance can be maintained between the rotor blade tips and the housing.

[0040] The metal bushing primarily serves as an airflow path surface. In one embodiment, the metal bushing can be a passive system that relies on the movement of blades to generate airflow. In another embodiment, the metal bushing can be additionally driven by the exhaust flow to fill the cavity between the inner diameter surfaces of the metal bushing and the composite housing. For example, a relatively high-temperature exhaust flow from the compressor outlet can circulate around the metal bushing to warm and expand the metal bushing, which is cooler than the warmer exhaust flow. As a result, the temperature of the warmer exhaust flow decreases simultaneously due to contact with the cooler metal bushing. The cooler exhaust flow can then be used to provide cooling to one or more components in the engine's turbine that are exposed to much hotter combustion gases.

[0041] In one embodiment, the metal bushing is spring-loaded onto the inner radial surface of the housing to enable concentric alignment with the blades. In another embodiment, the metal bushing is radially positioned above the tips of the rotor blades. In yet another embodiment, spring arms that can be integrated with or detached from the metal bushing may be provided. These spring arms may be segmented and spaced circumferentially around the housing, extending 360 degrees around the longitudinal centerline of the turbine engine. The metal bushing may include an optional layer of wear-resistant material. The metal bushing may have a substantially constant thickness, be tapered, or may include reinforced rear ribs or arms. In yet another embodiment, the metal bushing may include an optional exhaust gas flow opening to supply airflow into a cavity between the metal bushing and the inner radial surfaces of the housing. In yet another embodiment, a spacer material (e.g., a negative CTE material) may be provided between the metal bushing and the housing. When a positive thermal gradient is applied, the metal bushing expands while the negative CTE material spacer contracts to maintain a substantially constant and relatively small gap between the rotor blade tips and the metal bushing. The rotor blades also expand due to (i) the centrifugal force generated by the high rotational speed of the rotor blades and (ii) the relatively high temperature of the airflow flowing around the rotor blades.

[0042] Now refer to the attached diagram, Figure 1 This is a schematic cross-sectional view of a turbine engine 100 that can be used on an aircraft according to an embodiment of the present disclosure. The turbine engine 100 has an axial direction A (extending parallel to a longitudinal centerline (axis) 101, the longitudinal centerline (axis) 101 in... Figure 1 (Shown for reference) the radial direction R and the circumferential direction C. The circumferential direction C extends in the direction of rotation about the longitudinal centerline (axis) 101 (axial direction A). Figure 1 In the depicted embodiment, the turbine engine 100 is a non-ducted fan engine or an open fan engine. The turbine engine 100 has three different thrust-generating airflows during operation. Figure 1 The "three-flow engines" (labeled S1, S2, and S3) are described in further detail below. The turbocharger 100 includes a fan section 102 and a turbocharger 104 disposed downstream of the fan section 102.

[0043] Figure 1 The turbocharged engine 104 depicted includes a compressor section 110, a combustion section 120, and a turbine section 130 in a series flow relationship. The turbocharged engine 104 is substantially surrounded within a core shroud 106, which is substantially tubular and annularly surrounds the turbocharged engine 104. The core shroud 106 defines a core inlet 141, and in this embodiment, the core inlet 141 is annular. Figure 1 As schematically shown, compressor section 110 includes a turbocharger or low-pressure (LP) compressor 112, followed downstream by a high-pressure (HP) compressor 114. Combustion section 120 is downstream of compressor section 110. Turbine section 130 is downstream of combustion section 120 and includes a high-pressure (HP) turbine 132, followed downstream by a low-pressure (LP) turbine 134. Turbocharged engine 104 also includes a core air exhaust nozzle 143 (also referred to as an injection exhaust nozzle) downstream of turbine section 130. Compressor section 110, combustion section 120, and turbine section 130 together at least partially define a core airflow path (also referred to as a core duct 140) extending from core inlet 141 to core air exhaust nozzle 143, through which core air A5 flows. As will be discussed in more detail below, turbocharged engine 104 includes a high-pressure (HP) shaft 108 or HP spool and a low-pressure (LP) shaft 109. HP shaft 108 drives HP turbine 132 to HP compressor 114. HP turbine 132 and HP compressor 114 rotate synchronously via HP shaft 108. LP shaft 109 drives LP turbine 134 to LP compressor 112. LP turbine 134 and LP compressor 112 rotate synchronously via LP shaft 109.

[0044] Each of the LP compressor 112 and HP compressor 114 may include multiple compressor stages. In each stage, multiple compressor blades 116 rotate relative to corresponding multiple static compressor impellers 118 (also referred to as nozzles) to compress or pressurize the core air A5 through that stage. In a single compressor stage, the multiple compressor blades 116 may be arranged in a ring extending radially outward from the blade platform relative to a longitudinal centerline (axis) 101 to the blade tip (e.g., extending in the radial direction R). The compressor blades 116 may be part of a compressor rotor including a disk, wherein the compressor blades 116 extend radially from the disk. Other configurations of the compressor rotor may be used, including, for example, impeller disks, wherein the disk and the compressor blades 116 are integrally formed as a single piece. The corresponding static compressor impellers 118 are positioned upstream of and adjacent to the rotating compressor blades 116. The compressor impellers 118 of a stage of the compressor may be mounted circumferentially to the core housing 107. The core housing 107 may at least partially define a core airflow path (core duct 140). Each compressor stage can be used to sequentially compress core air A5 flowing through the core airflow path (core duct 140) to generate compressed air A6. Any suitable number of compressor blades 116, compressor impellers 118, and compressor stages can be used.

[0045] Each of the HP turbine 132 and LP turbine 134 may further include multiple turbine stages. In each stage, multiple turbine blades 136 rotate relative to corresponding multiple static turbine blades 138 (also referred to as nozzles) to extract energy from the combustion gas A7 passing through the stage. The turbine blades 136 may be part of a turbine rotor. Any suitable configuration of the turbine rotor may be used, including, for example, a disk from which the multiple turbine blades 136 extend. The corresponding static turbine blades 138 are positioned upstream of and adjacent to the rotating turbine blades 136. The turbine blades 138 of the first stage of the turbine may be mounted to the core housing 107 in a circumferential arrangement.

[0046] In combustion section 120, fuel received from the fuel system (not shown) is injected into combustion chamber 124 of combustor 122 via fuel nozzle 126. The fuel is mixed with compressed air A6 from compressor section 110 to form a fuel-air mixture, which is then burned to produce combustion products (i.e., combustion gases A7). As will be discussed further below, adjusting the fuel metering unit (not shown) of the fuel system alters the amount of fuel supplied to combustion chamber 124, and thus changes the amount of propulsive thrust generated by turbine engine 100. Combustion gases A7 are discharged from combustion chamber 124. These combustion gases can be directed to turbine blades 136 of HP turbine 132, and then to turbine blades 136 of LP turbine 134, and the combustion gases A7 drive (rotate) the turbine blades 136 of both HP turbine 132 and LP turbine 134. Any suitable number of turbine blades 136, turbine wheel blades 138, and turbine stages can be used. After flowing through the turbine section 130, the combustion gas A7 is discharged from the turbine engine 100 through the core air exhaust nozzle 143 to provide propulsive thrust.

[0047] The turbocharged engine 100, and more specifically, the turbocharged engine 104, also includes one or more drive shafts. As described above, the turbocharged engine 104 includes a high-pressure (HP) shaft 108 drivingly connecting an HP turbine 132 to an HP compressor 114, and a low-pressure (LP) shaft 109 drivingly connecting an LP turbine 134 to an LP compressor 112. More specifically, the turbine rotor of the HP turbine 132 is connected to the HP shaft 108, and the compressor rotor of the HP compressor 114 is connected to the HP shaft 108. Combustion gas A7 is directed into and expands through the HP turbine 132, wherein a portion of the thermal or kinetic energy from the combustion gas A7 is extracted via one or more stages of turbine blades 136 and turbine wheel blades 138 of the HP turbine 132. This causes the HP shaft 108 to rotate, which supports the operation (self-sustaining cycle) of the HP compressor 114 and the rotation of the compressor rotor, thereby supporting the rotation of the compressor blades 116 of the HP compressor 114. In this manner, combustion gas A7 performs work on HP turbine 132. Combustion gas A7 is then directed into LP turbine 134 and expands through LP turbine 134. Here, a second portion of thermal or kinetic energy is extracted from combustion gas A7 via one or more stages of turbine blades 136 and turbine wheel blades 138 of LP turbine 134. This causes LP shaft 109 to rotate, which supports the operation (self-sustaining cycle) of LP compressor 112 and the rotation of the compressor rotor, thereby supporting the rotation of compressor blades 116 of LP compressor 112. In this manner, combustion gas A7 performs work on LP turbine 134. HP shaft 108 and LP shaft 109 are coaxially arranged about longitudinal centerline (axis) 101. The diameter of HP shaft 108 is larger than the diameter of LP shaft 109, and HP shaft 108 is located radially outside of LP shaft 109. HP shaft 108 and LP shaft 109 are rotatable about longitudinal centerline (axis) 101 and are coupled to rotatable elements, such as compressor rotors and turbine rotors, as discussed above.

[0048] Figure 1 The fan section 102 shown includes a fan, referred to herein as primary fan 150. In the depicted embodiment, primary fan 150 is an open rotor fan, also known as a non-ducted fan. Primary fan 150 has a plurality of primary fan blades 151 coupled to fan disk 153. Figure 1 As depicted, the primary fan blades 151 extend outward from the fan disk 153 in a generally radial direction R. Figure 1The image depicts a primary fan blade 151, but multiple primary fan blades 151 can be arranged at equal intervals around a longitudinal centerline (axis) 101. The primary fan blades 151 and the fan disk 153 can rotate together about the longitudinal centerline (axis) 101 via a fan shaft 155. The fan disk 153 is covered by a fan hub 157, which is aerodynamically shaped to facilitate airflow through the multiple primary fan blades 151. In this embodiment, the fan hub 157 can rotate together with the primary fan blades 151 and the fan disk 153.

[0049] like Figure 1 As shown, the fan shaft 155 is connected to the LP shaft 109 via a reduction gearbox or power gearbox (also referred to as gearbox assembly 159). Therefore, the LP shaft 109 is driven to the primary fan. The gearbox assembly 159 is... Figure 1 The diagram is schematically shown. Gearbox assembly 159 includes multiple gears for adjusting the rotational speed of fan shaft 155, thereby adjusting the rotational speed of primary fan 150 relative to LP shaft 109. Gearbox assembly 159 can be used to reduce the rotational speed to a speed more efficient for primary fan 150. Gearbox assembly 159 can have gear ratios of 4:1 to 14:1, or 7:1 to 12:1, or 4:1 to 10:1, or 5:1 to 9:1, or 6:1 to 9:1, and can be constructed in a planetary or rotary star configuration. For non-ducted fan engines (e.g., turbine engine 100), the gearbox assembly can have gear ratios of 4:1 to 14:1. The gearbox can be a single-stage gearbox or a compound gearbox (e.g., having multiple stages).

[0050] In the case of variable pitch fans, such as Figure 1 As depicted, for example, a plurality of primary fan blades 151 are rotatable relative to the fan disk 153 about a fan blade pitch axis PB. Each of the primary fan blades 151 can be connected to the fan disk 153 via a pitch bearing 162, which allows rotation of the primary fan blade 151 about the pitch axis PB. The primary fan blade 151 can be rotated within the pitch bearing 162 by a pitch actuator 164, which is operatively coupled to the primary fan blade 152 to change the pitch of the corresponding primary fan blade 151. One or more pitch actuators 164 can be used, and in some embodiments, the pitch actuators 162 cause the primary fan blades 151 to rotate uniformly. A fan actuation system 160 controls one or more pitch actuators 164 to change the pitch of the primary fan blades 151 about their respective pitch axes PB. The fan actuation system 160 can be disposed within the fan hub 157.

[0051] Fan section 102 includes multiple fan guide vanes 172 ( Figure 1(Only one is shown in the image). Fan guide vanes 172 are circumferentially spaced and arranged around the longitudinal centerline (axis) 101 as part of the fan guide vane array 170. Figure 1 In the depicted embodiment, the fan guide vane 172 is a static airfoil and cannot rotate about the longitudinal centerline (axis) 101. Each fan guide vane 172 is mounted to the fan shroud 174 and extends outward from the fan shroud 174 generally along the radial direction R. In the case of a variable pitch fan, as... Figure 1 As depicted, for example, multiple fan guide vanes 172 are rotatable relative to the fan shroud 174 about the fan guide vane pitch axis PV. A fan guide vane actuation system 166 can be used to change the pitch of the fan guide vanes 172. The fan guide vane actuation system 166 can operate similarly to the fan actuation system 160 discussed above, and that discussion applies here. Therefore, the same reference numerals are used for the pitch bearing 162 and the pitch actuator 164 of the fan guide vane actuation system 166.

[0052] A fan shroud 174 annularly surrounds at least a portion of the core shroud 106 and is positioned generally along the radial direction R outside the core shroud 106. The fan shroud 174 and the core shroud 106 together define the casing of the turbine engine 100. A downstream section of the fan shroud 174 extends above the front portion of the core shroud 106 to define a fan flow path, also referred to as a fan duct 181. Incoming air enters through the fan duct inlet 183, passes through the fan duct 181, and exits through the fan exhaust nozzle 185 to generate propulsive thrust. The fan duct 181 is an annular duct positioned generally along the radial direction R outside the core duct 140. The fan shroud 174 and the core shroud 106 are connected together and are supported by a plurality of struts 176 ( Figure 1 (Only one support is shown in the image.) The struts 176 are circumferentially spaced around the longitudinal centerline (axis) 101 and extend radially outward from the core shroud 106. Each of the multiple struts 176 has an aerodynamic profile to guide the airflow therefrom.

[0053] The turbine engine 100 includes an inlet duct 187. The inlet duct 187 extends between an engine inlet 189, a core inlet 141, and a fan duct inlet 183. The engine inlet 189 is generally defined at the front end of a fan shroud 174 and positioned along the axial direction A between a primary fan 150 and a fan guide vane 172. The inlet duct 187 is an annular duct positioned along the radial direction R inside the fan shroud 174. Air flowing downstream along the inlet duct 187 is split (but not necessarily uniformly) by a splitter 178 of the core shroud 106 to the core duct 140 and the fan duct 181. The inlet duct 187 is wider along the radial direction R than the core duct 140. The inlet duct 187 is also wider along the radial direction R than the fan duct 181.

[0054] Fan section 102 also includes an intermediate fan 190. The intermediate fan 190 includes multiple intermediate fan blades 192. Figure 1 (Only one is shown). A plurality of intermediate fan blades 192 are rotatable about a longitudinal centerline axis 101. In the depicted embodiment, the intermediate fan 190 is drivenly coupled to the LP turbine 134 via an LP shaft 109. The plurality of intermediate fan blades 192 may be arranged at equal circumferential spacing about the longitudinal centerline axis 101. The intermediate fan blades 192 may be part of a rotor including a central hub (such as a disk 194), and each of the plurality of intermediate fan blades 192 extends radially from the disk 194. Other configurations of the compressor rotor may be used, including, for example, an impeller disk, wherein the disk 194 and the intermediate fan blades 192 are integrally formed as a single piece.

[0055] Multiple intermediate fan blades 192 are annularly surrounded by a fan shroud 174 (e.g., duct type). In this respect, an intermediate fan 190 is positioned inside the fan shroud 174 along the radial direction R. The intermediate fan 190 is positioned within an inlet duct 187 upstream of the core duct 140 and the fan duct 181. The ratio of the span of the primary fan blade 151 to the span of the intermediate fan blade 192 (span is measured from the root to the tip of the respective blade) can be selected as needed. Multiple inlet intermediate fan blades 196 are positioned upstream of and adjacent to the rotating intermediate fan blades 192. The inlet intermediate fan blades 196 can be mounted to the fan shroud 174 in a circumferential arrangement.

[0056] During operation of the turbine engine 100, the initial airflow, or incoming air A1, passes through the primary fan blades 151 of the primary fan 150 and splits into a first airflow (the first portion of air, referred to herein as primary bypass air A2) and a second airflow (the second portion of air, referred herein as engine air A3). The primary bypass air A2 bypasses the engine inlet 189 and flows radially R on the outside of the fan shroud 174, generally along the axial direction A. The primary bypass air A2 is accelerated by the primary fan blades 151 and passes through the fan guide vanes 172. The primary bypass air A2 then continues downstream to generate the primary propulsion flow, or first thrust flow S1. Most of the net thrust generated by the turbine engine 100 is produced by the first thrust flow S1.

[0057] Engine air A3 is directed or directed into inlet duct 187 and enters inlet duct 187 through engine inlet 189. Flowing downstream through inlet duct 187, engine air A3 flows over the intermediate fan blades 192 of intermediate fan 190 and is compressed by the rotating intermediate fan blades 192. After flowing over the intermediate fan blades 192, engine air A3 is split by splitter 178 into a third airflow (third portion air, referred herein as secondary bypass air A4) and a fourth airflow (fourth portion air, referred herein as core air A5). Core air A5 is directed or directed into the upstream section of core duct 140, or more specifically, into core inlet 141. Core air A5 flows through core duct 140 (as discussed above) to generate combustion gases A7 and exits core duct 140 through core air exhaust nozzle 143 to generate core airflow, also referred to as second thrust flow S2.

[0058] Secondary bypass air A4 is directed or directed into fan duct 181 and enters fan duct 181 through fan duct inlet 183. Secondary bypass air A4 flows through fan duct 181 generally in the axial direction A and exits fan duct 181 through fan duct inlet 183 to generate a third flow, also referred to as the third thrust flow S3. The third thrust flow S3 is a small fraction of the secondary airflow that increases fluid energy to generate total propulsion system thrust. In some embodiments, the pressure ratio of the third flow is higher than that of the primary propulsion flow (e.g., bypass or propeller-driven propulsion flow). Thrust can be generated by a dedicated nozzle or by mixing the secondary airflow with the primary propulsion flow or core airflow (e.g., into a common nozzle). In some embodiments, the operating temperature of the secondary airflow is lower than the engine's maximum compressor discharge temperature. Furthermore, in some embodiments, aspects of the third flow (e.g., airflow characteristics, mixing characteristics, or exhaust characteristics), and thus its percentage contribution to total thrust, are passively adjusted during engine operation, or can be purposefully modified using engine control features (such as fuel flow, motor power, variable stator, variable inlet guide vanes, valves, variable exhaust geometry, or fluid characteristics) to adjust or improve overall system performance across a wide range of potential operating conditions.

[0059] Figure 1The turbine engine 100 shown and discussed herein (e.g., a non-ducted fan engine) is provided by way of example only. In other embodiments, any other suitable engine may be used with aspects of this disclosure. For example, in other embodiments, the engine may be any other suitable gas turbine engine, such as a high-bypass turbofan engine, a turboshaft engine, a turboprop engine, a turbojet engine, etc. In this way, in other embodiments, the gas turbine engine may have other suitable configurations, such as other suitable numbers or arrangements of shafts, compressors, turbines, fans, etc. Furthermore, although turbine engine 100 is shown as a geared variable-pitch turbofan engine, in other embodiments, turbine engine 100 may be a direct-drive turbine engine or may be a fixed-pitch turbine engine. In other embodiments, the primary fan 150 and the fan guide vane array 170 may be ducted or shielded, including a nacelle or shield that circumferentially surrounds one or both of the primary fan 150 and the fan guide vane array 170, either jointly or individually. Furthermore, in alternative embodiments, aspects of this disclosure may be incorporated into or otherwise used with any other type of engine (such as a reciprocating engine).

[0060] The turbine engine 100 discussed herein is suitable for use on aircraft. Suitable aircraft include, for example, airplanes and unmanned aerial vehicles (UAVs). In other embodiments, the turbine engine can be any other turbine engine, such as an industrial turbine engine incorporated into a power generation system, or a marine turbine engine on a ship or other vessel.

[0061] Figure 2 This is a cross-sectional view of a portion of the inlet duct 187 including the intermediate fan 190 according to an embodiment of the present disclosure, showing... Figure 1 Detail 2. In the following paragraphs, the inlet duct 187 of the intermediate fan 190, including the intermediate fan blades 192 (rotor blades), is used only as an example to illustrate various features of this disclosure. The application of this disclosure is not limited to the inlet duct 187 of the intermediate fan 190 including the intermediate fan blades 192, but can be applied to any area of ​​the turbine engine 100 using rotor blades and housing. As described above, the fan shroud 174 ( Figure 1 (As shown) partially defines the inlet pipe 187 and circumferentially surrounds the central fan blade 192. Figure 2(Only one blade is shown in the image). More specifically, the fan shroud 174 may include one or more housings or housing segments to define the inlet duct 187. In the depicted embodiment, the fan housing 200 is positioned radially outside the intermediate fan blade 192 of the intermediate fan 190. The fan housing 200 is located radially to surround the intermediate fan blade 192 in a surrounding relationship. Each of the intermediate fan blades 192 includes a tip 198. In an embodiment, the fan housing 200 is made of a composite material.

[0062] In one embodiment, the fan housing 200 includes an annular housing body 201 and a metal bushing 220, the metal bushing 220 being positioned adjacent to the inner radial surface 212 of the annular housing body 201 of the fan housing 200. In another embodiment, the metal bushing 220 is disposed between the tip 198 of the intermediate fan blade 192 and the inner radial surface 212 of the annular housing body 201 of the fan housing 200. The metal bushing 220 is separable from the annular housing body 201 of the fan housing 200. The tip 198 of the intermediate fan blade 192 is spaced apart from the metal bushing 220 of the housing 200 by a gap G along the radial direction R. Because the annular housing body 201 of the fan housing 200 is made of a composite material, the annular housing body 201 of the fan housing 200 does not have the same thermal expansion or contraction as the metallic material of the intermediate fan blade 192. Initially, a gap G with an initial size D is provided between the intermediate fan blade 192 and the annular housing body 201 of the fan housing 200. However, the size D of the clearance G may increase due to the difference in thermal expansion or contraction between the intermediate fan blade 192 and the annular housing body 201. The efficiency of the turbine engine depends at least in part on the gap or clearance G between the fan housing 200 and the intermediate fan blade 192. If the size of the clearance G is too large, an excessive portion of the engine airflow will pass through the gap instead of interacting with the rotating intermediate fan blade 192, resulting in reduced engine efficiency. If the clearance is too small, interference may occur between the intermediate fan blade 192 and the fan housing 200, leading to damage to one or more surfaces of the intermediate fan blade 192 or the fan housing 200. As a result, a metal bushing 220 is used to substantially match the thermal expansion or contraction of the intermediate fan blade 192 so as to maintain a smaller clearance G with a tighter gap between the intermediate fan blade 192 and the fan housing 200, thereby not reducing engine efficiency or damaging one or more surfaces of the intermediate fan blade 192 or the fan housing 200.

[0063] The coefficient of thermal expansion (CTE) of the annular housing body 201 is lower than that of the metal bushing 220 and the intermediate fan blade 192. The metal bushing 220 is provided to control the gap between the metal bushing 220 and the tip 198 of the intermediate fan blade 192. In an embodiment, the metal bushing 220 may be made of the same metallic material as the intermediate fan blade 192 (e.g., titanium or a titanium alloy). In an embodiment, the metal bushing 220 includes a frame 220A and a plurality of spring arms 220B extending from the frame 220A. In an embodiment, as... Figure 2 As shown, the spring arm 220B is integrally formed of the same material as the frame 220A. In another embodiment, the spring arm 220B may be separate from the frame 220A (e.g., attached to the frame 220A). In one embodiment, a plurality of spring arms 220B may be segmented and spaced apart around the fan housing 200 in the circumferential direction C. In another embodiment, a plurality of spring arms 220B may be positioned around the turbine engine 100 along the circumferential direction C. Figure 1 The longitudinal centerline axis 101 (as shown in the diagram) extends the entire circumference (360 degrees) of the fan housing 200. In an embodiment, the metal bushing 220 may have a substantially constant thickness T along the axial direction A (e.g., ...). Figure 2 (as shown in the diagram). In another embodiment, the metal bushing 220 may also have a tapered thickness T that follows the profile of the tip 198 of the intermediate fan blade 192.

[0064] In an embodiment, such as Figure 2 As shown, one or more exhaust airflow openings 200A can be provided in the annular housing body 201 of the fan housing 200 to supply airflow 200B to a cavity 222 defined by the frame 220A of the metal bushing 220 and the inner radial surface 212 of the annular housing body 201 of the fan housing 200. For example, airflow 200B at a relatively high temperature from the compressor discharge port can circulate within the cavity 222 defined by the frame 220A of the metal bushing 220, causing the metal bushing 220, which is at a lower temperature than the warm airflow 200B, to warm and expand. As a result, the temperature of the warmer airflow 200B supplied through the one or more exhaust airflow openings 200A decreases simultaneously due to contact with the cooler metal bushing 220. The cooler airflow exiting through the one or more exhaust airflow openings 200A can then be used to supply airflow to the turbine engine 10 ( Figure 1 One or more components (as shown) exposed to combustion gases at much higher temperatures provide a cooling airflow. The cooling airflow can be provided inside one or more components, or through orifices in one or more components. For example, the cooling airflow can provide an air-cooled layer between one or more components and the combustion gases.

[0065] The metal bushing 220 may also include an abrasive surface 220C. The abrasive surface 220C of the metal bushing 220 is configured to be abraded by the tip 198 of the intermediate fan blade 192 when it encounters the abrasive surface 220C. In an embodiment, the abrasive surface 220C may be a layer of abrasive material (e.g., ceramic material) deposited on the frame 220A of the metal bushing 220.

[0066] In one embodiment, the metal bushing 220 may abut against the inner radial surface 212 of the annular housing body 201 of the fan housing 200, such that the plurality of spring arms 220B of the metal bushing 220 abut against the inner radial surface 212 of the annular housing body 201. The frame 220A of the metal bushing 220 may have a U-shaped form, such as... Figure 2 As shown in the diagram. However, the frame 220A of the metal bushing 220 can also have other shapes, such as a trapezoidal shape. Figure 2 As shown, the frame 220A of the metal bushing 220 is fixed in place at one end along the axial direction A by a notch or lip 212A on the inner radial surface 212, and is fixed in place at the opposite end along the axial direction A by an axial retaining member 214.

[0067] During operation, when a thermal gradient is present, such as during the operation of the turbine engine 100, the multiple spring arms 220B can expand under a positive thermal gradient (i.e., an increase in temperature) or retract under a negative thermal gradient (i.e., a decrease in temperature). As a result, the multiple spring arms 220B will exert a force on or retract from the inner radial surface 212 of the annular housing body 201 of the fan housing 200, which in turn maintains the gap G between the fan housing 200 and the tip 198 of the intermediate fan blade 192 within the desired tight clearance.

[0068] In an embodiment, the metal bushing 220 can be used as an airflow path surface. For example, the engine airflow path is at the intermediate fan blade 192 ( Figure 1 (as shown in the diagram) passes between. The tip 198 of the intermediate fan blade 192 rotates very close to the wearable surface 220C (inner surface) of the metal bushing 220. The wearable surface 220C of the metal bushing 220 also corresponds to the passage through the turbine engine 10 (shown in the diagram). Figure 1 The turbine engine 10 with intermediate fan 190 (as shown) Figure 1 The outer surface of the flow path (shown in the diagram). In one embodiment, the metal bushing 220 may be a passive system that relies on the movement of the blades 192 to generate airflow. In another embodiment, the metal bushing 220 may additionally be driven by an airflow 200B flowing through an exhaust airflow opening 200A to fill the cavity 222 between the metal bushing 220 of the fan housing 200 and the inner radial surface 212 of the annular housing body 201.

[0069] Figure 3 This is a cross-sectional view of component 300 taken along the radial direction R according to an embodiment of the present disclosure. Component 300 includes a fan housing 301 having an annular housing body 303, a metal bushing 302, and a plurality of metal spring extensions 304. For example, component 300 can be located in inlet pipe 187 ( Figure 1 (as shown in the illustration). In the depicted embodiment, component 300, including fan housing 301, is positioned at the intermediate fan blade 192 (as illustrated in the illustration). Figure 3 Only one is shown in the image. The metal bushing 302 is radially spaced from the annular housing body 303 of the fan housing 301. A plurality of metal spring extensions 304 are disposed between the annular housing body 303 and the metal bushing 302. In one embodiment, the annular housing body 303 may be made of a composite material. In one embodiment, the plurality of metal spring extensions 304 are coupled to the metal bushing 302. In another embodiment, the plurality of metal spring extensions 304 may be integrally formed with the metal bushing 302 and may be made of the same metal (e.g., titanium or a titanium alloy). In one embodiment, the metal spring extensions 304 contact the inner radial surface 301A of the annular housing body 303. In one embodiment, the plurality of metal spring extensions 304 may be segmented and spaced apart around the circumferential direction C of the annular housing body 303. In one embodiment, as... Figure 3 As shown, multiple metal spring extensions 304 form an angle relative to the circumference of the metal bushing 302. In an embodiment, as... Figure 3 As shown, the plurality of metal spring extensions 304 are oriented counterclockwise. However, the plurality of metal spring extensions 304 may also be oriented clockwise. The plurality of metal spring extensions 304 may be arranged to be axially aligned with the metal bushing 302, or may be oriented in a helical manner.

[0070] In this embodiment, a metal bushing 302 is disposed between the tip 198 of the intermediate fan blade 192 and the inner radial surface 301A of the annular housing body 303 of the fan housing 301. The tip 198 of the intermediate fan blade 192 is spaced apart from the fan housing 301 (e.g., the metal bushing 220 of the fan housing 301) by a gap G along the radial direction R. Since the annular housing body 303 of the fan housing 301 is made of a composite material that does not have the same thermal expansion as the metal material of the intermediate fan blade 192, the metal bushing 302 of the fan housing 301 is used to substantially match the thermal expansion of the intermediate fan blade 192, thereby maintaining a tighter gap G between the tip 198 of the intermediate fan blade 192 and the assembly 300.

[0071] In one embodiment, the metal bushing 302 may be made of the same metallic material as the intermediate fan blade 192 (e.g., titanium or a titanium alloy). In another embodiment, the metal bushing 302 may have a substantially constant thickness T along the axial direction A. In yet another embodiment, the metal bushing 302 may also have a tapered thickness T that follows the profile of the tip 198 of the intermediate fan blade 192.

[0072] The metal bushing 302 may include a wearable surface 306. The wearable surface 306 of the metal bushing 302 is configured to be worn by the tip 198 of the intermediate fan blade 192 when it encounters the wearable surface 306. In an embodiment, the wearable surface 306 may be a layer of wearable material (e.g., ceramic material) deposited on the metal bushing 302.

[0073] During operation, when a positive thermal gradient (i.e., temperature rise) is present, such as during the operation of the turbine engine 100, the metal bushing 302 and the plurality of metal spring extensions 304 can expand and lengthen under the positive thermal gradient. As a result, the plurality of metal spring extensions 304 apply a force to the inner radial surface 301A of the annular housing body 303 of the fan housing 301. Consequently, the metal bushing 302 in the first direction (in Figure 3 In the illustrated configuration, the fan blades rotate clockwise circumferentially and move radially toward the annular housing body 303 of the fan housing 301 to control the gap G between the intermediate fan blades 192 and the fan housing 301 (e.g., the metal bushing 302 of the fan housing 301). When a positive thermal gradient is present, the intermediate fan blades 192 also expand. Therefore, the radial movement of the metal bushing 302 toward the annular housing body 303 provides the ability to counteract the expansion of the intermediate fan blades 192, thus maintaining a relatively constant gap.

[0074] On the other hand, when a negative thermal gradient exists (i.e., a decrease in temperature), the metal bushing 302 and the plurality of metal spring extensions 304 can retract and collapse under the thermal gradient. As a result, the plurality of metal spring extensions 304 release the force on the inner radial surface 301A of the annular housing body 303 of the fan housing 301. As a result, the metal bushing 302 in the second direction opposite to the first direction (i.e., in the direction opposite to the first direction) Figure 3 In the configuration shown, the annular housing body 303 rotates counterclockwise and moves radially away from the fan housing 301 to control the gap G between the intermediate fan blade 192 and the assembly 300, and more specifically, to control the gap G between the intermediate fan blade 192 and the fan housing 301 (e.g., the metal bushing 302 of the fan housing 301).

[0075] Figure 4A It is according to an embodiment of the present disclosure along the axial direction A in Figure 3An axial view of the metal bushing 302 taken at line 4A-4A. Figure 4A A metal spring extension 304 is shown, oriented at an angle relative to the surface of the metal bushing 302. The metal spring extension 304 is arranged in a helical configuration similar to the teeth of a helical gear, in which the teeth are set at an angle relative to the axis or rotation.

[0076] Figure 4B This is an axial view of a metal bushing 302 along axial direction A according to another embodiment of the present disclosure. In this embodiment, the metal spring extension 304 is in contact with... Figure 4A The orientation shown is positioned relative to each other in the direction.

[0077] Figure 4C This is an axial view of a metal bushing 302 along axial direction A according to another embodiment of the present disclosure. In this embodiment, the metal spring extension 304 is oriented in a direction normal to the surface metal bushing 302.

[0078] Figure 5 According to embodiments of this disclosure Figure 3 A cross-sectional view of the metal bushing 302 taken at line 5-5 in the diagram. (See image.) Figure 5 As shown, the tip 198 of the intermediate fan blade 192 may have an angled profile relative to the leading edge 192A and the trailing edge 192B of the intermediate fan blade 192. The metal bushing 302 and its optional wearable surface 306 may be configured such that the metal bushing 302 and the wearable surface 306 follow the angled profile of the tip 198 of the intermediate fan blade 192. A clearance G is provided between the tip 198 of the intermediate fan blade 192 and the wearable surface 306 of the metal bushing or metal bushing 302.

[0079] Figure 6 This is a cross-sectional view of component 600 taken along the radial direction R according to an embodiment of the present disclosure. Component 600 includes a fan housing 601 having a metal bushing 602, an annular housing body 603, and a plurality of spacers 604. For example, component 600 may be located in inlet pipe 187 ( Figure 1 (as shown in the illustration). In the depicted embodiment, component 600, including fan housing 601, is positioned within intermediate fan blade 192 (as shown in the illustration). Figure 6(Only one is shown in the image) radially outward. The metal bushing 602 is radially spaced from the fan housing 601. A plurality of spacers 604 are disposed between the annular housing body 603 of the fan housing 601 and the metal bushing 602 of the fan housing 601. In an embodiment, the annular housing body 603 of the fan housing 601 may be made of a composite material. In an embodiment, the plurality of spacers 604 are made of a material with negative thermal expansion (i.e., a negative CTE material). A negative CTE material is a material that expands as the temperature decreases (negative thermal gradient) and contracts as the temperature increases (positive thermal gradient). An example of a negative CTE material is ALLVAR alloy manufactured by ALLVAR Corporation, Texas. ALLVAR alloy is a titanium-based alloy.

[0080] In one embodiment, a plurality of spacers 604 are distributed around the circumference of the metal bushing 602 and the circumference of the annular housing body 603 of the fan housing 601. In another embodiment, the plurality of spacers 604 contact the inner radial surface 601A of the annular housing body 603 of the fan housing 601 and the outer radial surface 602A of the metal bushing 602.

[0081] In this embodiment, a metal bushing 602 is disposed between the tip 198 of the intermediate fan blade 192 and the inner radial surface 601A of the annular housing body 603 of the fan housing 601. The tip 198 of the intermediate fan blade 192 is spaced apart from the metal bushing 602 by a gap G along the radial direction R. Since the annular housing body 603 of the fan housing 601 is made of a composite material, the annular housing body 603 does not have the same thermal expansion as the metallic material of the intermediate fan blade 192. Furthermore, the metal bushing 602 is used to substantially match the thermal expansion of the intermediate fan blade 192, thereby maintaining the gap G with a tighter clearance.

[0082] In an embodiment, the variation in the radial length L of the plurality of spacers 604 can be matched to the expansion or contraction of the radius R1 of the metal bushing 602. For example, when the radius R1 of the metal bushing 602 increases radially outward (i.e., expands) under a positive thermal gradient (i.e., temperature rises), the radial length L of the spacers 604 decreases (i.e., contracts), offsetting the increase in the radius R1 of the metal bushing 602, thereby reducing the stress between the metal bushing 602 and the annular housing body 603 (while the radius R2 of the annular housing body 603 remains substantially constant). Similarly, when the radius R1 of the metal bushing 602 decreases radially inward (i.e., retracts) under a negative thermal gradient (i.e., temperature falls), the radial length L of the spacers 604 increases (i.e., expands), offsetting the decrease in the radius R1 of the metal bushing 602, thereby reducing the stress between the metal bushing 602 and the annular housing body 603 (while the radius R2 of the annular housing body 603 remains substantially constant).

[0083] The term "substantially constant" is used herein to indicate that the variation in radius R2 of the annular housing body 603 of the fan housing 601 is negligible (e.g., less than 5%). When the radius R1 of the metal bushing 602 increases or decreases, the thickness T of the metal bushing 602 may increase or decrease, or may not increase or decrease.

[0084] In one embodiment, the metal bushing 602 may be made of the same metal material as the intermediate fan blade 192 (e.g., titanium or a titanium alloy). In another embodiment, the metal bushing 602 may have a substantially constant thickness along the axial direction A. In yet another embodiment, the metal bushing 602 may also have a tapered thickness that follows the profile of the tip 198 of the intermediate fan blade 192.

[0085] The metal bushing 302 may include a wearable surface 606. The wearable surface 606 of the metal bushing 602 is configured to be worn by the tip 198 of the intermediate fan blade 192 when it encounters the wearable surface 606. In an embodiment, the wearable surface 606 may be a layer of wearable material (e.g., ceramic material) deposited on the metal bushing 602.

[0086] Figure 7 This is a cross-sectional view of the metal bushing 602 according to an embodiment of the present disclosure. Figure 7 As shown, a plurality of spacers 604 may be disposed between the metal bushing 602 and the fan housing 601, and disposed in the recess 700 of the fan housing 601.

[0087] This disclosure provides a blade clearance control system tailored for a turbine engine blade disk system, wherein the CTE of the housing is lower than that of the rotor system (e.g., the blade). Blades are typically made of metal (such as titanium), while the housing is made of a composite material. Using a composite housing instead of a metal housing allows for weight reduction of the system. In the presence of a thermal gradient, blades tend to thermally expand, while the housing does not expand or expands only minimally. Due to the thermal expansion of the blades, a clearance is provided between the blade tip and the housing to accommodate this expansion. Initially, a clearance of size D is provided between the blade and the composite housing. However, the size D of the clearance may increase due to the difference in thermal expansion or contraction between the blade and the composite housing. The efficiency of a turbine engine depends at least in part on the clearance or void between the composite housing and the blade. If the clearance is too large, an excessive portion of the engine airflow will pass through the clearance instead of interacting with the rotating blade, resulting in reduced engine efficiency. If the clearance is too small, interference may occur between the blade and the composite housing, leading to damage to one or more surfaces of the blade or the composite housing. Therefore, metal bushings (with or without spacers) are used to substantially match the thermal expansion or contraction of the blades, maintaining a smaller clearance with a tighter gap, thus without reducing engine efficiency or damaging one or more surfaces of the blades or composite housing. Higher performance can be achieved by providing a smaller or tighter clearance between the blade tip and the housing. As a result, a more matched CTE system is provided, allowing the housing and blades to expand by similar amounts to maintain a tighter clearance.

[0088] Further details are provided by the following topics.

[0089] An assembly for controlling the clearance between rotor blades and a housing in a turbine engine, the assembly including a housing comprising an annular housing body and a bushing positioned adjacent to an inner radial surface of the annular housing body, the housing being radially spaced from the rotor blades by the clearance. The coefficient of thermal expansion (CTE) of the annular housing body is lower than the CTE of the bushing and the CTE of the rotor blades, and the bushing is configured to control the clearance between the rotor blades and the housing.

[0090] According to the components described in the foregoing clauses, the housing includes a composite housing, the bushing includes a metal bushing, and the rotor blades include metal rotor blades.

[0091] According to any of the preceding clauses, the metal bushing is configured to match the thermal expansion or contraction of the metal rotor blades to maintain a smaller gap between the metal rotor blades and the composite housing with a tighter clearance, thereby not reducing engine efficiency or damaging one or more surfaces of the metal rotor blades or the composite housing. The metal bushing includes a frame and a plurality of spring arms extending from the frame, the plurality of spring arms abutting an inner radial surface of the annular housing body, the plurality of spring arms being configured to expand and apply force to the inner radial surface of the annular housing body, or retract from the inner radial surface of the annular housing body to maintain the gap.

[0092] According to any of the foregoing clauses, the plurality of spring arms are integrally formed of the same metal material as the frame.

[0093] According to any of the foregoing clauses, the plurality of spring arms are segmented and spaced apart in the circumferential direction around the composite housing.

[0094] According to any of the foregoing clauses, the plurality of spring arms extend over the entire 360-degree circumference of the composite housing.

[0095] According to any of the foregoing clauses, the metal bushing has a substantially constant thickness along the axial direction of the turbine engine.

[0096] According to any of the foregoing clauses, the frame has a U-shaped form, and the plurality of spring arms are disposed in the internal cavity of the U-shaped form.

[0097] According to any of the preceding clauses, the frame of the metal bushing is fixed in place at one end along the axial direction of the turbine engine by a notch in the inner radial surface and fixed in place at the opposite end along the axial direction by an axial retaining member.

[0098] According to any of the foregoing clauses, the metal bushing has a tapered thickness that follows the profile of the tip of the rotor blade.

[0099] According to any of the preceding clauses, the composite housing includes one or more exhaust airflow openings to supply air to a cavity defined by the inner radial surface of the frame of the metal bushing and the annular housing body.

[0100] According to any of the preceding clauses, the metal bushing includes a wearable surface configured to be worn by the tips of the rotor blades, and the wearable surface is a layer of wearable material deposited on the frame of the metal bushing.

[0101] The component according to any of the foregoing clauses further includes a plurality of spring extensions disposed between the annular housing body and the bushing, the plurality of spring extensions contacting an inner radial surface of the housing, wherein the plurality of spring extensions are configured to (i) expand and lengthen under a positive thermal gradient to apply a force to the inner radial surface of the housing, thereby circumferentially rotating the bushing in a first direction and radially moving the bushing toward the housing to control the gap between the rotor blades and the housing; and (ii) retract and shrink under a negative thermal gradient to release the force on the inner radial surface of the housing, thereby circumferentially rotating the bushing in a second direction and radially moving the bushing away from the housing to control the gap between the rotor blades and the housing, wherein the annular housing body comprises a composite annular housing body, the bushing comprises a metal bushing, and the plurality of spring extensions comprises a plurality of metal spring extensions, wherein the plurality of metal spring extensions are integrally formed with the metal bushing.

[0102] According to any of the foregoing clauses, the plurality of metal spring extensions are segmented and spaced circumferentially around the composite annular housing body, and the plurality of metal spring extensions form an angle relative to the circumference of the metal bushing.

[0103] According to any of the foregoing clauses, the metal bushing is made of the same metal material as the rotor blades.

[0104] According to any of the foregoing clauses, the component wherein the metal bushing includes a wearable surface configured to be worn by the tips of the rotor blades.

[0105] According to any of the foregoing clauses, the metal bushing has a tapered thickness that follows the profile of the tip of the rotor blade.

[0106] The component according to any of the foregoing clauses further includes a plurality of spacers disposed between the annular housing body and the bushing, the plurality of spacers contacting both the inner radial surface of the annular housing body and the bushing, wherein the plurality of spacers are made of a material having a negative coefficient of thermal expansion (CTE), wherein when the radius of the bushing increases under a positive thermal gradient, the radial length of the plurality of spacers decreases to offset the increase in the radius of the bushing, and when the radius of the bushing decreases under a negative thermal gradient, the radial length of the plurality of spacers increases to offset the decrease in the radius of the bushing, thereby reducing the stress between the bushing and the annular housing body.

[0107] According to any of the foregoing clauses, the plurality of spacers are distributed around the circumference of the bushing and the circumference of the housing.

[0108] According to any of the preceding clauses, the bushing includes a wearable surface, the wearable surface of the bushing being worn by the tip of the rotor blade when the tip of the rotor blade encounters the wearable surface of the bushing.

[0109] A turbine engine includes an assembly for controlling a clearance between rotor blades and a housing in the turbine engine. The assembly includes a housing comprising an annular housing body and a bushing positioned adjacent to an inner radial surface of the annular housing body, the housing being radially spaced from the rotor blades by the clearance. The coefficient of thermal expansion (CTE) of the annular housing body is lower than the CTE of the bushing and the CTE of the rotor blades, and the bushing is configured to control the clearance between the rotor blades and the housing.

[0110] According to the turbine engine described in the foregoing clause, the housing includes a composite housing, the bushing includes a metal bushing, and the rotor blades include metal rotor blades.

[0111] According to any of the preceding clauses of the turbine engine, the metal bushing is configured to match the thermal expansion or contraction of the metal rotor blades to maintain a smaller clearance between the metal rotor blades and the composite housing with a tighter gap, thereby not reducing engine efficiency or damaging one or more surfaces of the metal rotor blades or the composite housing. The metal bushing includes a frame and a plurality of spring arms extending from the frame, the plurality of spring arms abutting an inner radial surface of the annular housing body, the plurality of spring arms being configured to expand and apply force to the inner radial surface of the annular housing body, or retract from the inner radial surface of the annular housing body to maintain the clearance.

[0112] The turbine engine according to any of the foregoing clauses, wherein the plurality of spring arms are integrally formed of the same metal material as the frame.

[0113] The turbine engine according to any of the foregoing clauses, wherein the plurality of spring arms are segmented and spaced apart in the circumferential direction around the composite housing.

[0114] The turbine engine according to any of the foregoing clauses, wherein the plurality of spring arms extend the entire 360-degree circumference of the composite housing.

[0115] The turbine engine according to any of the foregoing clauses, wherein the metal bushing has a substantially constant thickness along the axial direction of the turbine engine.

[0116] The turbine engine according to any of the foregoing clauses, wherein the frame has a U-shaped shape and the plurality of spring arms are disposed in the internal cavity of the U-shaped shape.

[0117] According to any of the preceding clauses, the frame of the metal bushing is fixed in place at one end along the axial direction of the turbine engine by a notch in the inner radial surface and fixed in place at the opposite end along the axial direction by an axial retaining member.

[0118] The turbine engine according to any of the foregoing clauses, wherein the metal bushing has a tapered thickness that follows the profile of the tip of the rotor blade.

[0119] According to any of the preceding clauses, the turbine engine, wherein the composite housing includes one or more exhaust airflow openings to supply air to a cavity defined by the inner radial surface of the frame of the metal bushing and the annular housing body.

[0120] According to any of the preceding clauses, the turbine engine wherein the metal bushing includes a wearable surface configured to be worn by the tips of the rotor blades, and the wearable surface is a layer of wearable material deposited on the frame of the metal bushing.

[0121] The turbine engine according to any of the foregoing clauses further includes a plurality of spring extensions disposed between the annular housing body and the bushing, the plurality of spring extensions contacting an inner radial surface of the housing, wherein the plurality of spring extensions are configured to (i) expand and lengthen under a positive thermal gradient to apply a force to the inner radial surface of the housing, thereby circumferentially rotating the bushing in a first direction and radially moving the bushing toward the housing to control the clearance between the rotor blades and the housing; and (ii) retract and shrink under a negative thermal gradient to release the force on the inner radial surface of the housing, thereby circumferentially rotating the bushing in a second direction and radially moving the bushing away from the housing to control the clearance between the rotor blades and the housing, wherein the annular housing body comprises a composite annular housing body, the bushing comprises a metal bushing, and the plurality of spring extensions comprises a plurality of metal spring extensions, wherein the plurality of metal spring extensions are integrally formed with the metal bushing.

[0122] According to any of the preceding clauses, the plurality of metal spring extensions are segmented and spaced circumferentially around the composite annular housing body, and the plurality of metal spring extensions form an angle relative to the circumference of the metal bushing.

[0123] In any of the preceding clauses, the turbine engine wherein the metal bushing is made of the same metal material as the rotor blades.

[0124] The turbine engine according to any of the foregoing clauses, wherein the metal bushing includes a wearable surface configured to be worn by the tips of the rotor blades.

[0125] The turbine engine according to any of the foregoing clauses, wherein the metal bushing has a tapered thickness that follows the profile of the tip of the rotor blade.

[0126] The turbine engine according to any of the foregoing clauses further includes a plurality of spacers disposed between the annular housing body and the bushing, the plurality of spacers contacting both the inner radial surface of the annular housing body and the bushing, wherein the plurality of spacers are made of a material having a negative coefficient of thermal expansion (CTE), wherein when the radius of the bushing increases under a positive thermal gradient, the radial length of the plurality of spacers decreases to offset the increase in the radius of the bushing, and when the radius of the bushing decreases under a negative thermal gradient, the radial length of the plurality of spacers increases to offset the decrease in the radius of the bushing, thereby reducing the stress between the bushing and the annular housing body.

[0127] The turbine engine according to any of the foregoing clauses, wherein the plurality of spacers are distributed around the circumference of the bushing and the circumference of the housing.

[0128] According to any of the preceding clauses of the turbine engine, wherein the bushing includes a wearable surface, the wearable surface of the bushing being worn by the tip of the rotor blade when the tip of the rotor blade encounters the wearable surface of the bushing.

[0129] According to any of the foregoing clauses, the turbine engine wherein the plurality of metal spring extensions are oriented counterclockwise or clockwise.

[0130] The turbine engine according to any of the foregoing clauses, wherein the metal bushing has a substantially constant thickness along the axial direction.

[0131] The turbine engine according to any of the foregoing clauses, wherein the metal bushing has a tapered thickness that follows the profile of the tip of the rotor blade.

[0132] The turbine engine according to any of the foregoing clauses, wherein the first direction is clockwise and the second direction is counterclockwise.

[0133] The turbine engine according to any of the foregoing clauses, wherein the material having the negative coefficient of thermal expansion includes titanium-based alloys.

[0134] The turbine engine according to any of the foregoing clauses, wherein the plurality of spacers are distributed around the circumference of the bushing and the circumference of the housing.

[0135] The turbine engine according to any of the foregoing clauses, wherein the plurality of spacers are in contact with the inner radial surface of the housing and with the outer radial surface of the bushing.

[0136] According to any of the preceding clauses of the turbine engine, wherein the bushing has a wearable surface, the wearable surface of the bushing being worn by the tip of the rotor blade when the tip of the rotor blade encounters the wearable surface of the bushing.

[0137] Although the foregoing description is directed to preferred embodiments of the present disclosure, other variations and modifications will be apparent to those skilled in the art and can be made without departing from the present disclosure. Furthermore, features described in connection with one embodiment of the present disclosure may be used in conjunction with other embodiments, even if not explicitly stated above.

Claims

1. An assembly to control a clearance between a rotor blade and a casing in a turbine engine, characterized by, The components include: The housing includes an annular housing body and a bushing positioned adjacent to the inner radial surface of the annular housing body, the housing being separated from the rotor blades in the radial direction by a gap. The coefficient of thermal expansion (CTE) of the annular housing body is lower than that of the bushing and the rotor blades, and the bushing is configured to control the gap between the rotor blades and the housing.

2. The assembly of claim 1, wherein, The device further includes a plurality of spacers disposed between the annular housing body and the bushing, the plurality of spacers contacting both the inner radial surface of the annular housing body and the bushing, wherein the plurality of spacers are made of a material having a negative coefficient of thermal expansion (CTE). Specifically, when the radius of the bushing increases under a positive thermal gradient, the radial length of the plurality of spacers decreases to offset the increase in the radius of the bushing, and when the radius of the bushing decreases under a negative thermal gradient, the radial length of the plurality of spacers increases to offset the decrease in the radius of the bushing, thereby reducing the stress between the bushing and the annular housing body.

3. The assembly of claim 2, wherein, in, The plurality of spacers are distributed around the circumference of the bushing and the circumference of the housing.

4. The assembly of claim 2, wherein, in, The bushing includes a wearable surface, which is configured to be worn by the tip of the rotor blade when the tip of the rotor blade encounters the wearable surface of the bushing.

5. The assembly of claim 1, wherein, The device further includes a plurality of spring extensions disposed between the annular housing body and the bushing, the plurality of spring extensions contacting the inner radial surface of the housing. The plurality of spring extensions are configured to (i) expand and lengthen under a positive thermal gradient to apply a force to the inner radial surface of the housing, thereby circumferentially rotating the bushing in a first direction and radially moving the bushing toward the housing to control the gap between the rotor blades and the housing, and (ii) retract and shrink under a negative thermal gradient to release the force on the inner radial surface of the housing, thereby circumferentially rotating the bushing in a second direction and radially moving the bushing away from the housing to control the gap between the rotor blades and the housing. The annular housing body comprises a composite annular housing body, the bushing comprises a metal bushing, and the plurality of spring extensions comprises a plurality of metal spring extensions. The plurality of metal spring extensions are integrally formed with the metal bushing.

6. The assembly of claim 5, wherein, in, The plurality of metal spring extensions are segmented and spaced apart circumferentially around the composite annular housing body, and the plurality of metal spring extensions form an angle relative to the circumference of the metal bushing.

7. The assembly of claim 5, wherein, in, The metal bushing is made of the same metal material as the rotor blades.

8. The assembly of claim 5, wherein, in, The metal bushing includes a wearable surface configured to be worn by the tips of the rotor blades.

9. The component according to claim 5, characterized in that, in, The metal bushing has a tapered thickness that follows the profile of the tip of the rotor blade.

10. The assembly of claim 1, wherein, in, The annular housing body comprises a composite annular housing body, the bushing comprises a metal bushing, and the rotor blade comprises a metal rotor blade.

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