Gas turbine engine including a rotating blade assembly

Abstract

A rotating blade assembly for a turbine engine having a drive shaft includes a disk, at least one blade assembly, and a retainer. The disk is operably coupled to the drive shaft and includes a seat having at least a portion of a first through hole. The at least one blade assembly has an upper platform, a lower platform, a dovetail extending from the lower platform, and a blade extending between the upper platform and the lower platform. The retainer assembly secures the disk to the at least one blade assembly and includes a hollow tubular element, a pin, and a fastener.

Description

[0001] Cross-reference to related applications

[0002] This application claims priority to Italian patent application No. 102021000029963, filed on November 26, 2021, the entire contents of which are incorporated herein by reference. Technical Field

[0003] This disclosure generally relates to gas turbine engines, and more specifically, to rotating blade assemblies of gas turbine engines. Background Technology

[0004] A turbine engine, particularly a gas turbine engine, is a rotating engine that extracts energy from a flow of working air that passes in series through a compressor section (where working air is compressed), a combustor section (where fuel is added to the working air and ignited), and a turbine section (where the combusted working air expands and does work to drive the compressor section and other systems), and provides thrust in an aircraft implementation. The compressor stage and turbine stage comprise axially arranged pairs of rotating blades and stationary blades. A gas turbine engine may be arranged as an engine core comprising at least a compressor section, a combustor section, and a turbine section arranged in an axial flow configuration, and defining at least one rotating element or rotor and at least one stationary component or stator. Attached Figure Description

[0005] The complete and enabling disclosure of this description, including its best mode, is set forth in the specification with reference to the accompanying drawings, for those skilled in the art, wherein:

[0006] Figure 1 This is a schematic cross-sectional view of a turbine engine for an aircraft comprising a counter-rotating turbine section and a counter-rotating compressor section, based on various aspects described herein.

[0007] Figure 2 From Figure 1 Section II Figure 1 A cross-sectional view of the counter-rotating turbine section of the turbine engine, further showing the rotating blade assembly including the disk, blade assembly and retainer assembly.

[0008] Figure 3 yes Figure 2 An exploded perspective view of the rotating blade assembly further shows the disk, blade assembly, and retainer assembly.

[0009] Figure 4 From Figure 3 The cross-sectional view of the rotating blade assembly seen in section IV-IV further illustrates the retainer assembly.

[0010] Figure 5 is a cross-sectional view of an exemplary rotating blade assembly of a gas turbine engine of Figure 1 , further illustrating an exemplary disk and an exemplary blade assembly.

[0011] Figure 6 is a radial view of a rotating blade assembly of Figure 5 , further illustrating a tail of an exemplary disk housed within the exemplary blade assembly.

[0012] Figure 7 is a cross-sectional view of an exemplary rotating blade assembly of a gas turbine engine of Figure 1 , further illustrating an exemplary disk and an exemplary retainer assembly including a retainer plate.

[0013] Figure 8 is a cross-sectional view of an exemplary rotating blade assembly of a gas turbine engine of Figure 1 , further illustrating an exemplary disk and an exemplary retainer assembly including a retainer plate.

[0014] Figure 9 is a schematic axial view of an exemplary rotating blade assembly of a gas turbine engine of Figure 1 , further including an exemplary disk and a set of retainers disposed along a portion of the disk corresponding to each other blade assembly. DETAILED DESCRIPTION

[0015] Aspects of the present disclosure relate to a rotating blade assembly for a gas turbine engine including a drive shaft. The rotating blade assembly further includes a disk operably coupled to the drive shaft and including a seat having at least a portion of a first bore; and at least one blade assembly having a dovetail. A retainer assembly can secure the at least one blade assembly to the disk. Aspects of the present disclosure are described in particular to gas turbine engines, including reverse rotating sections. In other words, reverse rotating gas turbine engines. As used herein, the term “reverse rotating section” or iterations thereof can refer to a portion of a gas turbine engine including a set of axially adjacent, serially arranged rotating components (e.g., blades) that rotate in opposite circumferential directions. However, it will be appreciated that although described in the context of reverse rotating gas turbine engines, aspects of the present disclosure described herein are not so limited and can have universal applicability in any suitable geared gas turbine engine, turboprop engine, turboshaft engine, or turbofan engine in non-limiting examples. However, it will be further appreciated that aspects of the present disclosure described herein are not so limited and can have universal applicability in other gas turbine engines. For example, the present disclosure can be applicable to rotating blade assemblies in other engines or vehicles and can be used to provide benefits in industrial, commercial, and residential applications.

[0016] As used herein, the terms “forward” or “upstream” refer to movement in a direction toward an inlet of the gas turbine engine, or a component that is relatively closer to the inlet of the gas turbine engine as compared to another component. The terms “aft” or “downstream” used in connection with “forward” or “upstream” refer to a direction toward an aft or outlet of the gas turbine engine, or a component that is relatively closer to the outlet of the gas turbine engine as compared to another component.

[0017] As used herein, “a set” can include any number of individually described elements, including only one element. Further, the terms “radial” or “radially” as used herein refer to a dimension extending between a central longitudinal axis of the gas turbine engine and an outer circumference of the engine.

[0018] All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, back, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, aft, etc.) are only used for identification purposes to aid the reader’s understanding of the present disclosure, and do not create limitations, particularly as to the position, orientation, or use of the disclosure. Connection references (e.g., attached, coupled, connected, and joined) are to be construed broadly and will be given their ordinary meanings to refer to the connection between two elements component including intermediate members between the elements, and relative movement between elements unless otherwise indicated. As such, connection references do not necessarily infer that two elements are directly connected and in fixed relation to each other. The exemplary drawings are for purposes of illustration only and the dimensions, positions, order and relative sizes reflected in the attached drawings attached hereto can vary.

[0019] Figure 1 is a schematic cross-sectional view of a gas turbine engine 10 for an aircraft. The gas turbine engine 10 has a generally longitudinally extending axis or centerline 12 extending from a forward direction 14 to an aft direction 16. The gas turbine engine 10 can include at least one counter-rotating section. Thus, the gas turbine engine 10 can be defined as a counter-rotating gas turbine engine. The gas turbine engine 10 includes, in downstream serial flow relationship, a fan section 18 including a forward fan assembly 20 and an aft fan assembly 21, a counter-rotating compressor section 22 including at least one counter-rotating section, a combustion section 28 including a combustor 30, a counter-rotating turbine section 32 including at least one counter-rotating section, and an exhaust section 38.

[0020] In the illustrated gas turbine engine 10, the counter-rotating compressor section 22 can include a counter-rotating low pressure (LP) compressor 24 and a counter-rotating high pressure (HP) compressor 26, while the counter-rotating turbine section 32 can include a counter-rotating HP turbine 34 and a counter-rotating LP turbine 36. It will be appreciated that aspects of the present disclosure can be applicable to other turbine engines, including engines without any counter-rotating sections, or turbine engines without counter-rotating portions. By way of non-limiting example, aspects of the present disclosure can be applicable to other gas turbine engines that do not include a counter-rotating LP turbine. For example, a turbine engine is contemplated having a LP turbine with a static circumferentially arranged vane axially spaced apart from a rotating circumferentially arranged blade.

[0021] As shown, fan assemblies 20 and 21 are positioned at a forward end of gas turbine engine 10. The terms "forward fan" and "aft fan" are used herein to indicate that one of fan assemblies 20 is coupled axially upstream of aft fan assembly 21. It is also contemplated that fan assemblies 20, 21 can be positioned at an aft end of gas turbine engine 10. Fan assemblies 20 and 21 each include a plurality of fan blades 40 positioned within a fan casing 42. Fan blades 40 are joined to respective rotor disks 44 that are rotatably coupled to forward fan assembly 20 by a respective forward fan shaft 46 and to aft fan assembly 21 by an aft fan shaft 47.

[0022] Reverse rotating HP compressor 26, combustor 30, and reverse rotating HP turbine 34 form an engine core 48 of gas turbine engine 10. Gas turbine engine core 48 is surrounded by an outer casing 50, which can be coupled with fan casing 42. Reverse rotating HP turbine 34 is coupled to reverse rotating HP compressor 26 via a core rotor or shaft 52. In operation, gas turbine engine core 48 produces combustion gases that are directed downstream to reverse rotating LP turbine 36, which extracts energy from the gases for powering fan assemblies 20, 21 through their respective fan shafts 46, 47.

[0023] Reverse rotating LP turbine 36 includes an outer rotor 54 positioned radially inward from outer casing 50. Outer rotor 54 can have a generally frustoconical shape and includes a first set of airfoils 56 arranged circumferentially and extending radially inward toward engine centerline 12.

[0024] Reverse rotating LP turbine 36 further includes an inner rotor 58 arranged substantially coaxially with respect to outer rotor 54 and radially inward of outer rotor 54. Inner rotor 58 includes a second set of airfoils 60 arranged circumferentially and axially spaced apart from first set of airfoils 56. Inner rotor 58 can be further defined as a first rotor, while outer rotor 54 can be defined as a second rotor. Second set of airfoils 60 extend radially outward away from engine centerline 12. First set of airfoils 56 and second set of airfoils 60 together define a plurality of turbine stages 62. In Figure 1 In the example shown, five turbine stages 62 are shown, and it should be appreciated that any number of stages can be used. Further, while first set of airfoils 56 is shown forward of second set of airfoils 60, first set of airfoils 56 and second set of airfoils 60 can be arranged in any suitable manner, including first set of airfoils 56 being positioned aft of second set of airfoils 60.

[0025] While the gas turbine engine 10 is described in the context of including a rotating outer rotor 54 and a rotating inner rotor 58, it is further contemplated that either the first set of airfoils 56 or the second set of airfoils 60 can be included in or form a portion of a stationary stator within the gas turbine engine 10. In one example, the first set of airfoils 56 can form a circumferentially arranged set of static vanes forming a portion of an outer stator within the gas turbine engine 10, while the second set of airfoils 60 are coupled to a rotatable inner rotor 58. In another example, the second set of airfoils 60 can be in the form of static vanes coupled to an inner stator within the gas turbine engine 10, where the first set of airfoils 56 can be in the form of blades coupled to an outer rotor.

[0026] In addition to the outer rotor 54 and the inner rotor 58, stationary portions of the gas turbine engine 10, such as the outer casing 50, are also individually or collectively referred to as a stator 63. Thus, the stator 63 can refer to the combination of non-rotating elements throughout the gas turbine engine 10.

[0027] In operation, the airflow exiting the fan section 18 is split such that a portion of the airflow is directed along the main flow path 15 into the counter-rotating LP compressor 24, which then supplies pressurized air 65 to the counter-rotating HP compressor 26, which further pressurizes the air. The pressurized air 65 from the counter-rotating HP compressor 26 is mixed with fuel in the combustor 30 and ignited, producing combustion gases 66 along the main flow path 15. Some work is extracted from these combustion gases 66 by the counter-rotating HP turbine 34, which drives the counter-rotating HP compressor 26. The combustion gases 66 exit along the main flow path 15 into the counter-rotating LP turbine 36, which extracts additional work to drive the counter-rotating LP compressor 24, and the exhaust gases ultimately exit the gas turbine engine 10 via the exhaust section 38. The driving of the counter-rotating LP turbine 36 can drive rotation of the forward fan assembly 20 and the counter-rotating LP compressor 24.

[0028] A portion of the pressurized air 65 can be extracted from the counter-rotating compressor section 22 as bleed air 67. The bleed air 67 can be extracted from the pressurized air 65 and provided to engine components that require cooling. The temperature of the pressurized air 65 entering the combustor 30 increases significantly to a temperature higher than that of the bleed air 67. The bleed air 67 can be used to reduce the temperature of core components downstream of the combustor.

[0029] Some air supplied by the fan 20, such as the bleed air 67, can bypass the gas turbine engine core 48 and be used to cool portions of the gas turbine engine 10, particularly hot portions of the gas turbine engine 10, or other portions of the gas turbine engine 10 or to power other portions of the gas turbine engine 10. In the context of turbine engines, the hot portions of the gas turbine engine are generally downstream of the combustor 30, particularly the counter-rotating turbine section 32, with the counter-rotating HP turbine 34 being the hottest portion as it is directly downstream of the combustion section 28. Other sources of cooling fluid can be, but are not limited to, fluid discharged from the counter-rotating LP compressor 24 or the counter-rotating HP compressor 26.

[0030] Figure 2 is seen from Figure 1 Figure 1 A cross-sectional axial view of the gas turbine engine 10 of FIG. 1 is shown. The gas turbine engine 10 can include a rotating vane assembly 100 within a portion of the gas turbine engine 10. The rotating vane assembly 100 can be disposed within a portion of the counter-rotating LP turbine 36. However, it should be appreciated that the rotating vane assembly 100 can be disposed within any suitable portion of the gas turbine engine 10, such as within any suitable portion of the counter-rotating compressor section 22 or the counter-rotating turbine section 32. Further, while a single rotating vane assembly 100 is shown, it should be appreciated that any number of one or more rotating vane assemblies 100 can be disposed within the gas turbine engine 10.

[0031] The rotating vane assembly 100 can include a vane assembly 102, a disk 104, and a set of retainer assemblies 106. At least a portion of the disk 104 can be operably coupled to a rotating component of the gas turbine engine 10. As a non-limiting example, the disk 104 can be operably coupled to the drive shaft 98 of the gas turbine engine 10. At least a portion of the vane assembly 102 can be operably coupled to another rotating component. As a non-limiting example, the vane assembly can be coupled to the inner rotor 58 or the outer rotor 54.

[0032] ​The vane assembly 102 can include an inner platform 108, an outer platform 110 located radially outward from the inner platform 108 relative to the engine centerline 12, and a set of circumferentially spaced apart vanes 112 extending therebetween. As shown, the set of circumferentially spaced apart vanes 112 can include the first set of airfoils 56. However, it should be understood that the set of circumferentially spaced apart vanes 112 can include any suitable set of airfoils, such as but not limited to the second set of airfoils 60. The set of circumferentially spaced apart vanes 112 can be any suitable vane or blade within the gas turbine engine 10 that is operably coupled to the outer rotor 54, the inner rotor 58, or a stationary portion (e.g., the stator 63) of the gas turbine engine 10. The outer platform 110 can be operably coupled to a rotating element of the gas turbine engine 10. As a non-limiting example, the outer platform 110 can be operably coupled to the inner rotor 58 or the outer rotor 54 of the gas turbine engine 10. The dovetail 114 can extend from a radially inner portion of the inner platform 108.

[0033] The disc 104 can extend between the outer periphery in the axial direction and the radial direction. The disc 104 can further surround or face at least a portion of the vane assembly 102. As a non-limiting example, the disc 104 can surround a radially inner portion of the vane assembly 102. As a non-limiting example, the disc 104 can surround a portion of the dovetail 114. The disc 104 can further include an inner periphery. At least a portion of the inner periphery can face or contact the vane assembly 102 or the retainer assembly 106. At least a portion of the inner periphery can define a seat 126 that at least partially surrounds or faces the dovetail 114. The seat 126 can be further defined by the first band 122 and the second band 124. As shown, the second band 124 can be disposed axially forward of or otherwise upstream of the first band 122 relative to the combustion gases 66. The disc 104 is sized and / or shaped such that the disc 104 can fit over or otherwise surround at least a corresponding portion of the dovetail 114.

[0034] The disk 104 can further include a protrusion 142 extending from a portion of the remainder of the disk 104. As shown, the protrusion 142 can extend from a portion of the first band 122. The protrusion 142 can be operatively coupled to the drive shaft 98 of the gas turbine engine 10. The drive shaft 98 can be any suitable drive shaft 98 as described herein, such as but not limited to, the forward fan shaft 46, the aft fan shaft 47, or the core shaft 52. Thus, with the rotating vane assembly 100 disposed within the counter-rotating turbine section 32, rotation of the rotating vane assembly 100 can be used to rotationally drive the drive shaft 98, which in turn, can drive an upstream portion of the gas turbine engine 10 (e.g., rotating components of the counter-rotating compressor section 22, a portion of the fan section 18, etc.). While the protrusion 142 is shown, it can be appreciated that the disk 104 can be formed without the protrusion 142. Thus, in some embodiments, the disk 104 is coupled to the rotating vane assembly 100 and not to a drive shaft.

[0035] The retainer assembly 106 can extend through a portion of the disk 104 and operatively couple the disk 104 to the vane assembly 102. As shown, the retainer assembly 106 can extend axially through the disk 104 and face axially opposite ends of the disk 104. Thus, the retainer assembly 106 can axially retain the disk 104 about a portion of the vane assembly 102. As a non-limiting example, the retainer assembly 106 can axially retain the disk 104 about the dovetail 114 of the vane assembly 102.

[0036] The vane assemblies 102 can be included in a set of circumferentially spaced apart vane assemblies 102. The set of vane assemblies 102 can extend around the entire engine centerline 12 to form a ring of vane assemblies 102. However, at least a portion of the disk 104 can extend continuously across the entire engine centerline 12. In other words, the disk 104 can form a 360 degree ring around the engine centerline 12. As such, the disk 104 can extend across one or more of the set of vane assemblies 102. Similarly, the protrusions 142 can be formed as a continuous ring or band that extends around the entire engine centerline 12. Alternatively, the protrusions 142 can be formed as separate segments such that the protrusions 142 are included in a set of segmented protrusions 142 that extend from respective portions of the disk 104. As such, the disk 104 can be formed as a hub and spoke assembly when viewed in a plane that is perpendicular to the engine centerline 12 and intersects the disk 104. The set of retainer assemblies 106 can include any suitable number of retainer assemblies 106 that are circumferentially spaced apart along the disk 104. As a non-limiting example, the set of retainer assemblies 106 can be regularly or otherwise equally circumferentially spaced apart around the disk 104. Alternatively, two or more of the retainer assemblies 106 can be formed in groups, where the retainer assemblies 106 are closer to each other than they are to the retainer assemblies 106 of adjacent groups. In any case, the disk 104 can be axially retained to each dovetail 114 of each vane assembly in the set of vane assemblies 102 via the set of retainer assemblies 106.

[0037] As shown, the vane assemblies 102, the disk 104, and the retainer assemblies 106 are separate components that are operably coupled to or otherwise face each other. However, it should be appreciated that at least a portion of the vane assemblies 102, the disk 104, or the retainer assemblies 106 can be integrally formed with another portion of the rotating vane assembly 100. As a non-limiting example, the retainer assemblies 106 can be integrally formed with the second band 124 or the first band 122 of the disk 104. As another non-limiting example, at least one of the first band 122 or the second band 124 can be integrally formed with the vane assemblies 102 such that the vane assemblies 102 and the disk 104 form a unitary structure.

[0038] Figure 3 is Figure 2 an exploded perspective view of the rotating vane assembly 100. The rotating vane assembly 100 can include the vane assemblies 102, the disk 104, and the retainer assemblies 106.

[0039] Each vane 112 of the set of circumferentially spaced apart vanes 112 can be defined by a leading edge 116 and a trailing edge 118, an outer wall downstream of or otherwise aft of the leading edge 116 in an axial direction, a root 119, and a tip 120. An extension of the outer wall between the leading edge 116 and the trailing edge 118 can define a chordwise direction. An extension of the outer wall between the root 119 and the tip 120 can define a spanwise direction. The root 119 can be coupled to or otherwise integrally formed with the inner platform 108. The tip 120 can be coupled to or otherwise integrally formed with the outer platform 110.

[0040] The first through-hole 128 can be defined by a portion of the disk 104. As a non-limiting example, the first band 122 and the second band 124 can each include a portion of the first through-hole 128. The radially inner portion or distal end of the dovetail 114 can further include a portion of the first through-hole 128. When assembled, the first through-hole 128 can continuously extend through the first band 122, the dovetail 114, and the second band 124. The first through-hole 128 can extend in the axial direction from an upstream portion to a downstream portion of the rotating vane assembly 100.

[0041] Each retainer assembly 106 of the set of retainer assemblies 106 can include a tubular element, a pin 132, and a fastener 134. The tubular element can include any suitable tubular shape when viewed in a plane that is perpendicular to the engine centerline 12 and intersects the tubular element, such as but not limited to a square tube, a cylindrical tube, or any other suitable tube. At least a portion of the retainer assembly 106 can face, contact, or be coupled to the seat 126 of the disk 104.

[0042] As a non-limiting example, the tubular element can be any suitable tubular element, such as the bushing 130. The bushing 130 can include a set of fingers 131 extending along a portion of the bushing 130. Each finger 131 of the set of fingers 131 can be separated from an adjacent finger 131 by a void or absence of material. The bushing 130 can include a hollow interior defining a second through-hole 136. Thus, the tubular element can be further defined as a hollow tubular element. The pin 132 can be aligned with and at least partially housed within the first through-hole 128 and the second through-hole 136. The pin 132 can terminate at a distal end 138. When assembled, the distal end 138 can extend beyond the first through-hole 128. The bushing 130 can be aligned with the first through-hole 128. The bushing 130 can further include a first end defining a shoulder 140 that, when assembled, can abut at least a portion of the disk 104. As a non-limiting example, the shoulder 140 can abut the second band 124. The fastener 134 can be secured to the distal end 138 of the pin 132 and abut a portion of the disk 104. The tubular element can generally be defined as any suitable element that can expand via a mechanical feature (e.g., the fingers 131) when an external force is applied to an interior of the tubular element. It is further contemplated that the tubular element can expand via a material property (e.g., thermal properties, elasticity, etc.). As a non-limiting example, the tubular element can be a rubber tube that expands when the pin 132 applies an external force.

[0043] Figure 4 is a cross-sectional view of the rotating vane assembly 100 as seen from a section IV-IV. Figure 3

[0044] The first band 122 can include a first rib 144. The second band 124 can include a second rib 146 opposite the first rib 144. As shown, the first rib 144 and the second rib 146 can be disposed on axially opposite sides of the dovetail 114. Both the first rib 144 and the second rib 146 can interface or otherwise contact corresponding portions of the dovetail 114. Both the first rib 144 and the second rib 146 can be used to radially retain the disk 104 on the vane assembly 102.

[0045] The second band 124 can include a portion that overlaps a corresponding portion of the first band 122. The overlapping portion can define a lap joint 145 formed between the first band 122 and the second band 124. The lap joint 145 can define a coupling or interface between the first band 122 and the second band 124. It is contemplated that the lap joint 145 can be further used to align the first band 122 relative to the second band 124.

[0046] ​As shown, the bushing 130 of the retainer assembly can extend through a portion of the first through hole 128 such that the second through hole 136 is aligned with the first through hole 128. It is contemplated that the bushing 130 can end along a distal end 148 of the bushing 130. The distal end 148 of the bushing 130 can be disposed within a portion of the first through hole 128. As a non-limiting example, the distal end 148 can be spaced apart from the first band 122 such that the bushing 130 is not in physical contact with the first band 122.

[0047] It is further contemplated that the first through hole 128 and the second through hole 136 can each include a non-constant cross-sectional area when viewed in a plane that is perpendicular to the engine centerline 12 and intersects the disk 104. As a non-limiting example, the first through hole 128 can include a region having a reduced cross-sectional area within a portion of the first band 122 when compared to the remainder of the first through hole 128. The portion of the first through hole 128 having the reduced cross-sectional area can directly contact at least a portion of the pin 132. As a non-limiting example, the bushing 130 can include a portion having a reduced cross-sectional area. In other words, the bushing 130 can include a portion in which the cross-sectional area linearly or non-linearly decreases. As a non-limiting example, at least a portion of the cross-sectional area of the bushing 130 can decrease from an upstream or axially forward portion to a downstream or axially rearward portion. Similarly, the pin 132 can include a reduced cross-sectional area when viewed in a plane that is perpendicular to the engine centerline 12 and intersects the pin 132. Thus, the pin 132 can be defined as a tapered pin and the bushing can be defined as a tapered bushing. The reduced cross-sectional area of the pin 132 can correspond to the reduced cross-sectional area of the bushing 130 such that the pin 132 can interface with the bushing 130 along a cross-section of the bushing 130 and pin 132 defined by the reduced cross-sectional area. The interface between the bushing 130 and the pin 132 can function to retain the pin 132 within the bushing 130.

[0048] The shoulder 140 of the bushing 130 can interface with the disk 104. As a non-limiting example, the shoulder 140 of the bushing 130 can interface with the second band 124. The second band 124 can include a cutout 150 that is sized to accommodate the shoulder 140 of the bushing 130. Thus, the forward portion of the bushing 130 can be flush with the forward portion of the second band 124.

[0049] It is contemplated that at least a portion of the fastener 134 can abut a portion of the first band 122. The fastener 134 can be any suitable fastener 134 such as, but not limited to, a nut, a hydraulic fastener, a magnetic fastener, a weld, an adhesive, or an electrical connection (e.g., an electrically actuated connection). As a non-limiting example, the fastener 134 can be defined by a nut. Thus, the distal end 138 can further include a threaded portion that corresponds to a threaded portion of the nut. Thus, the nut can be fastened or otherwise threadably connected to a threaded portion of the pin 132.

[0050] During assembly, the disk 104 can fit over the corresponding portion of the dovetail 114 such that the first through-hole 128 is formed continuously through the first band 122, the second band 124, and the dovetail 114. The first rib 144 and the second rib 146 can each interface with the corresponding portion of the dovetail 114. The size and positioning of the lap joint 145 can be designed to ensure that the first rib 144 and the second rib 146 are positioned within the correct position upon assembly. Further, the size of the lap joint 145 can then be designed to ensure that once the disk 104 is positioned over the dovetail 114, the first through-hole 128 is formed continuously through the disk 104 and the dovetail 114. The bushing 130 can then be aligned with and inserted into the first through-hole 128. The bushing 130 can be inserted such that the shoulder 140 contacts or is housed within the cutout 150. The pin 132 can then be inserted into the second through-hole 136 defined by the bushing 130. As discussed herein, the distal end 138 of the pin 132 can extend beyond the end of the first through-hole 128. The fastener 134 can then be placed, applied, fastened, or otherwise coupled to the distal end 138 of the pin 132 and abut a portion of the disk 104 (e.g., the first band 122). The fastener 134 can exert an axial fastening or closing force on the pin 132 to draw the pin 132 toward the fastener 134, which is axially constrained by the disk 104. As the pin 132 is drawn toward the fastener 134, the pin 132 first moves axially until it comes into contact with the bushing 130, where the pin 132 continues to move axially which then causes the shoulder 140 to abut the cutout 150 formed within the disk 104. With the shoulder 140 abutting the cutout 150, and the fastener 134 abutting the disk 104, opposing closing forces are exerted on opposite axial ends of the disk 104, which in turn, axially holds the disk 104 over the dovetail 114. Any additional axial movement of the pin 132 causes the fingers 131 of the bushing 130 to radially expand and exert an expanding clamping force between the disk 104 and the dovetail 114. Thus, the bushing 130 can be further defined as an expandable bushing or an expandable tubular element, respectively, it being understood that the expansion of the bushing 130 can be produced via any suitable method, such as a mechanical component, or material properties of the bushing 130. Similarly, the pin 132 can be generally defined as a component configured to actuate and expand the bushing 130 as described herein. The expanding clamping force, in turn, urges the dovetail 114 against the first rib 144 and the second rib 146. Thus, with this type of connection, the retainer assembly 106 is able to both axially constrain the disk 104 to the dovetail 114 and radially constrain the disk 104 to the dovetail 114.

[0051] During operation of the gas turbine engine 10, a working airflow can flow through a portion of the rotating vane assembly 100. As non-limiting examples, the working airflow can flow through a set of circumferentially spaced apart vanes 112 of the rotating vane assembly 100. As non-limiting examples, the working airflow can be any suitable airflow within the gas turbine engine, such as but not limited to the pressurized air 65 or the combustion gases 66. In instances where the rotating vane assembly 100 is disposed within the counter-rotating turbine section 32, the rotating vane assembly 100 can extract work from the working airflow as the working airflow flows through the rotating vane assembly 100. In instances where the rotating vane assembly 100 is disposed within the counter-rotating compressor section 22, the rotating vane assembly 100 can pressurize or otherwise compress the working airflow.

[0052] As discussed herein, the rotating vane assembly 100 can include a set of circumferentially spaced apart vane assemblies 102 that are separated from one another. It is contemplated that the disk 104 can be used to couple each vane assembly 102 of the set of vane assemblies 102 such that a continuous ring of vane assemblies 102 is formed. Thus, the rotating vane assembly 100 can be formed into a rigid rotating vane assembly 100 by the use of the disk 104 interconnecting the vane assemblies 102. This in turn can ensure that there is no or otherwise minimal radial clearance between the outer rotor 54 and the outer band 110 and between a portion of the disk 104 (e.g., the protrusion 142) and the drive shaft 98. The reduction or elimination of clearance can ensure that the rotation of the set of vanes 102 within the rotating vane assembly 100 is concentric with the rotation of the outer band 110 or the outer rotor 54. Similarly, the reduction or elimination of clearance can ensure that the disk 104 is concentric with the outer band 110. With the concentric rotation and assembly of the rotating vane assembly 100, the overall amount of losses is reduced (e.g., by the frictional losses of adjacent pieces abutting one another). This ultimately ensures that the overall efficiency of the gas turbine engine 10 is improved as compared to a gas turbine engine 10 that does not include the rotating vane assembly 100 with the disk 104.

[0053] During normal operation of the gas turbine engine 10, an operating force can be applied to the rotating vane assembly 100. The operating force can be defined as any suitable force (e.g., rotational force or thermal load) applied to the rotating vane assembly 100 during normal operation of the gas turbine engine 10. As a non-limiting example, a 30 klb operating force can be applied to the rotating vane assembly 100. The operating force can be applied to the disk 104 and define a radially inward force relative to the engine centerline 12. It is contemplated that the disk 104 can be formed to withstand these operating forces of the gas turbine engine 10. As a non-limiting example, the interface (e.g., the first and second ribs 144, 146) of the disk 104 with the dovetail 114 is sized or formed to withstand these operating forces. However, during shutdown of the gas turbine engine 10, a shutdown force, which is opposite the operating force, is applied to the rotating vane assembly 100. In other words, a radially outward force can be applied to the rotating vane assembly 100 during shutdown of the gas turbine engine 10. The shutdown force can be less than the operating force. As a non-limiting example, the shutdown force can be 0.05 times the operating force. As a non-limiting example, if the operating force is 30 klb, the shutdown force can be 1.5 klb. However, unlike the operating force, the shutdown force is transmitted through a portion of the retainer assembly 106 and not just the disk 104. Accordingly, the retainer assembly 106 is sized and formed to withstand the shutdown force. Because the shutdown force is much less than the operating force, the retainer assembly 106 can be formed from a weaker material than the disk 104, which ultimately reduces the material costs associated with the rotating vane assembly 100.

[0054] Figure 5 is an exemplary rotating vane assembly 200 within a gas turbine engine for Figure 1 FIG. 2 is a cross-sectional view of an exemplary rotating vane assembly 200 within a gas turbine engine. The exemplary rotating vane assembly 200 is similar to the rotating vane assembly 100; therefore, like components will be denoted with like numbers in the 200 series, and it should be understood that the description of the like components of the rotating vane assembly 100 applies to the exemplary rotating vane assembly 200 unless otherwise noted.

[0055] The rotating vane assembly 200 can include a vane assembly 202, a disk 204, and a retainer assembly 206, similar to the rotating vane assembly 100. The vane assembly 202, similar to the vane assembly 102, can include a vane 212 extending from a root 219 to a tip (not shown) and from a leading edge 216 to a trailing edge 218. The root 219 can be coupled to an inner platform 208 of the vane assembly 202, while the tip can be coupled to an outer platform (not shown) of the vane assembly 202. A dovetail 214 can depend from the inner platform 208. The disk 204, similar to the disk 104, can include a first band 222 and a second band 224 that together form a seat 226. The first band 222 can be operably coupled to a drive shaft, optionally through a tab 242. The retainer assembly 206, similar to the retainer assembly 106, can include a bushing 230, a pin 232, and a fastener 234. The bushing 230 can include a shoulder 240 and a set of fingers 231 that interface with a cutout 250 formed in a portion of the second band 224. The pin 232 can be defined by a distal end 238, and the fastener 234 can be secured to the distal end 238. At least a portion of the disk 204 and the dovetail 214 can form a continuous first through-hole 228. An interior of the bushing 230 can define a second through-hole 236 that is aligned with the first through-hole 228. The pin 232 can be at least partially disposed within the second through-hole 236 and the first through-hole 228. A lap joint 245 can be formed between the first band 222 and the second band 224 and define an interface or coupling between the first band 222 and the second band 224.

[0056] The disk 204, similar to the disk 104, includes a first band 222 and a second band 224. However, the first band 222 forms a plate that abuts the dovetail 214 and a portion of the second band 224. However, the second band 224 can be formed to extend only through a radially inner portion of the rotating vane assembly 200. In other words, the second band 224 does not extend forward of the dovetail 214 or otherwise face a portion of the dovetail 214 along a portion of the rotating vane assembly 200 that includes the retainer assembly 206. Further, the only portion of the first through-hole 228 defined by the second band 224 is the portion of the second band 224 that is opposite the dovetail 214. In other words, the second band 224 does not self-define all of the first through-hole 228 (e.g., the hole formed in the axially forward portion of the second band 124) as does the second band 12. The bushing 230 can extend through a portion of the first through-hole 228 and terminate within the first through-hole 228 at a terminal end 248.

[0057] Cutout 250 is similar to cutout 150; however, cutout 250 is also at least partially formed within dovetail 214. Therefore, the shoulder 240 of bushing 230 can abut at least a portion of disc 204 and dovetail 214 or blade assembly 202. As shown, the first through hole 228 has a smaller axial length than the first through hole 128. This is due to the construction of disc 204.

[0058] Bushing 230 is similar to bushing 130, except that bushing 230 may have a smaller axial length compared to bushing 130. This is due to the smaller axial length of the first through-hole 228. This, in turn, reduces the material required to manufacture the retainer assembly 206. Furthermore, bushing 230 may include a wall 251 terminating radially distally to define a shoulder 240.

[0059] During operation of the gas turbine engine 10, at least a portion of the working airflow can flow toward the disk 204, thus limiting leakage fluid. It is envisioned that minimizing leakage fluid within the gas turbine engine 10 could maximize the amount of working airflow flowing through the blades 212, which in turn maximizes the amount of work extracted from the working airflow. The bushing 230 can form a fluid seal between the disk 204 and the blade assembly 202 via the wall 251 and the shoulder 240. This, in turn, ensures that leakage fluid does not enter either the first through-hole 228 or the second through-hole 236. This reduces the total amount of leakage fluid, thereby maximizing the overall efficiency of the rotating blade assembly 200. It is envisioned that the remainder of the disk 204 can be used to limit leakage fluid. As a non-limiting example, the first band 222 can be used to reduce or otherwise eliminate leakage fluid that can flow from an upstream portion of the rotating blade assembly 200 to a downstream portion of the rotating blade assembly 200.

[0060] Figure 6 yes Figure 5 A radial view of the rotating blade assembly 200 as seen from a plane perpendicular to the engine centerline 12 and intersecting with the rotating blade assembly 200. The rotating blade assembly 200 may include blade assemblies 202, which are included within a set of blade assemblies 202 that are circumferentially spaced apart from each other.

[0061] Disc 204 can be connected to dovetail tenon 214 via a dovetail connector defined by tail 254 and socket 256. Disc 204 may include tail 254, which extends radially from the remainder of disc 204 relative to engine centerline 12. As a non-limiting example, second band 224 may include tail 254, which extends radially from the remainder of second band 224.

[0062] A socket 256 can be formed between circumferentially adjacent portions of adjacent vane assemblies 202. As shown, the socket 256 can be formed by circumferentially adjacent cuts formed within a portion of the dovetail 214. The socket 256 can be sized and shaped to accommodate the tail 254 of the disk 204.

[0063] During assembly of the rotatable vane assembly 200, the tail 254 of the disk 204 can be inserted through or into the socket 256. This can be accomplished by sliding the tail 254 and thus the disk 204 into the socket 256. At least a portion of the tail 254 can interface with the socket 256. The interface between the tail 254 and the socket 256 can radially retain the disk 204 to the vane assembly 202. This radial retention by the tail 254 and the socket 256 is similar to the radial retention between the first and second ribs 144, 146 and the dovetail 114 of the rotatable vane assembly 100. Further, at least a portion of the closing force that axially retains the disk 204 on the vane assembly 202 when the bushing 230 directly contacts the dovetail 214 can be applied directly to the vane assembly 202.

[0064] Figure 7 is Figure 1 A cross-sectional view of an exemplary rotatable vane assembly 300 of the gas turbine engine 10 of FIG. 1. The exemplary rotatable vane assembly 300 is similar to the rotatable vane assemblies 100, 200; therefore, like components will be denoted with like numbers in the 300 series, and it should be understood that the description of like components of the rotatable vane assemblies 100, 200 apply to the exemplary rotatable vane assembly 300 unless otherwise noted.

[0065] The rotating vane assembly 300 can include a vane assembly 302, a disk 304, and a retainer assembly 306, similar to the rotating vane assemblies 100, 200. The vane assembly 302, similar to the vane assemblies 102, 202, can include a vane 312 extending from a root 319 to a tip (not shown) and from a leading edge 316 to a trailing edge 318. The root 319 can be coupled to an inner platform 308 of the vane assembly 302, while the tip can be coupled to an outer platform (not shown) of the vane assembly 302. A dovetail 314 can depend from the inner platform 308. The disk 304, similar to the disks 104, 204, can define a seat 326. The disk 304 can be operably coupled to a drive shaft, optionally through a tab 342. The retainer assembly 306, similar to the retainer assemblies 106, 206, can include a bushing 330 having a set of fingers 331, a pin 332, and a fastener 334. The bushing 330 can be formed similarly to the bushing 230 in that it includes a wall 351 that terminates at a radially distal end to define a shoulder 340. The shoulder 340 can interface with a cutout 350 formed at least partially within a portion of the dovetail 314 and a portion of the disk 304. The bushing 330 can be further defined by a smaller axial length when compared to the bushing 130, similar to the bushing 230. The pin 332 can be defined by a distal end 338, and the fastener 334 can be secured to the distal end 338. At least a portion of the disk 304 and the dovetail 314 can form a continuous first through-hole 328. An interior of the bushing 330 can define a second through-hole 336 aligned with the first through-hole 328. The pin 332 can be at least partially disposed within the second through-hole 336 and the first through-hole 328. The bushing 330 can extend through a portion of the first through-hole 328 and terminate within the first through-hole 328 at a terminal end 348.

[0066] The disk 304 is similar to the disk 204 in that it is formed to extend only through a radially inner portion of the rotating vane assembly 200. In other words, the disk 204 does not contact or engage an axially forward portion of the dovetail 314. A difference between the disk 304 and the disk 204 is that the disk 204 includes the first band 222 and the second band 224. However, the disk 204 can be defined as a unitary disk 304 in which the first band 122, 222 is integrally formed with the second band 124, 224. In other words, the disk 304 is formed as a monolithic structure that is axially retained to the vane assembly 302 via the retainer assembly 306. The disk 304 can be further radially retained by using any suitable radial retention assembly as described herein (e.g., the tail 254 and the socket 256, or the first rib 144 and the second rib 146).

[0067] Figure 8 is Figure 1FIG. 4 is a cross-sectional view of an exemplary rotating blade assembly 400 of a gas turbine engine 10. The exemplary rotating blade assembly 400 is similar to the rotating blade assemblies 100, 200, 300; therefore, like components will be denoted with like numerals in the 400 series, and it should be understood that the description of like components of the rotating blade assemblies 100, 200, 300 apply to the exemplary rotating blade assembly 400 unless otherwise noted.

[0068] The rotating blade assembly 400 can include a blade assembly 402, a disk 404, and a retainer assembly 406, similar to the rotating blade assemblies 100, 200, 300. The blade assembly 402, similar to the blade assemblies 102, 202, 302, can include a blade 412 extending from a root 419 to a tip (not shown) and from a leading edge 416 to a trailing edge 418. The root 419 can be coupled to an inner platform 408 of the blade assembly 402, while the tip can be coupled to an outer platform (not shown) of the blade assembly 402. A dovetail 414 can depend from the inner platform 408. The disk 404, similar to the disks 104, 204, 304, can define a seat 426. The disk 404 can be operably coupled to a drive shaft, optionally through a protrusion 442. The disk 404 can be integrally formed disk 404, similar to the disk 304. The retainer assembly 406, similar to the retainer assemblies 106, 206, 306, can include a bushing 430, a pin 432, and a fastener 434. The bushing 430 can be integrally formed, similar to the bushings 130, 230, 330, in that it includes a shoulder 440 and a set of fingers 431. The pin 432 can be defined by a distal end 438, and the fastener 434 can be secured to the distal end 438. At least a portion of the disk 404 and the dovetail 414 can form a continuous first through-hole 428. An interior of the bushing 430 can define a second through-hole 436 aligned with the first through-hole 428. The pin 432 can be at least partially disposed within the second through-hole 436 and the first through-hole 428. The bushing 430 can extend through a portion of the first through-hole 428 and terminate within the first through-hole 428 at a terminal end 448.

[0069] The retainer assembly 406 can further include a retainer plate 458. The retainer plate 458 can abut the dovetail 414 and a portion of the disk 404. The retainer plate 458, similar to the disks 104, 204, 304 and the dovetails 214, 314, can include a cutout 450 configured to accommodate the shoulder 440 of the bushing 430. The retainer plate 458 can further include a through-hole 460 extending axially through a portion of the retainer plate 458. The retainer plate 458 can be aligned with the first through-hole 428 such that the through-hole 460 defines a portion of the first through-hole 428.

[0070] During assembly of the rotating vane assembly 400, at least the closing force or axial retention force generated by the retainer assembly 406 can be applied to the retainer plate 458. Thus, the retainer plate 458 can be used to axially retain the disk 404 to the vane assembly 402.

[0071] Figure 9 is Figure 1 a schematic axial view of an exemplary rotating vane assembly 500 of the gas turbine engine 10. The exemplary rotating vane assembly 500 is similar to the rotating vane assemblies 100, 200, 300, 400; thus, like parts will be denoted with like numbers in the 500 series, and it should be understood that the description of similar parts of the rotating vane assemblies 100, 200, 300, 400 apply to the exemplary rotating vane assembly 500 unless otherwise noted.

[0072] The rotating vane assembly 500 can include a set of vane assemblies 502, a disk 504, and a set of retainer assemblies 506, similar to the rotating vane assemblies 100, 200, 300, 400. Each vane assembly 502 of the set of vane assemblies 502 can be similar to the vane assemblies 102, 202, 302, 402. Each vane assembly 502 can include a vane 512 or a set of vanes 512, where each vane 512 extends from a root 519 to a tip (not shown) and from a leading edge 516 to a trailing edge (not shown). The root 519 can be coupled to an inner platform 508 of the vane assembly 502, while the tip can be coupled to an outer platform (not shown) of the vane assembly 502. Each vane assembly 502 can include a dovetail 514 that can overhang from a corresponding inner platform 508 of the vane assembly 502. As shown, the disk 504 is a schematic view of the disk 504. However, it should be understood that the disk 504 can be any suitable disk 104, 204, 304, 404 as disclosed herein. As shown, the set of retainer assemblies 506 is a schematic view of the retainer assemblies 506. However, it should be understood that each retainer assembly 506 of the set of retainer assemblies 506 can include any suitable retainer assembly 106, 206, 306, 406 as described herein.

[0073] As shown, each of the set of vane assemblies 502 can include a respective dovetail 514. Each dovetail 514 can be formed to be complementary to an adjacent dovetail 514. A first contact area 562 can be formed between corresponding portions of adjacent dovetails 514. The first contact area 562 can represent an area where adjacent dovetails 514 are in physical contact or otherwise coupled to one another. Further, each dovetail 514 can be in contact through a portion of the disc 504. As a non-limiting example, each dovetail 514 can be in contact through the disc 504 along a second contact area 564. As a non-limiting example, the second contact area 564 can be a contact or interface between the first rib 144 or the second rib 146 of the rotating vane assembly 100. As a non-limiting example, the second contact area 564 can be a contact or interface between the tail 254 and the socket 256 of the rotating vane assembly 200.

[0074] It is contemplated that the set of retainer assemblies 506 can be arranged in groups. As a non-limiting example, the set of retainer assemblies 506 can include a group of two retainer assemblies 506. However, it should be appreciated that each group of retainer assemblies 506 can include any number of one or more retainer assemblies 506. As shown, each group of retainer assemblies 506 of the set of retainer assemblies 506 can extend through a portion of the disc 504 corresponding to each other vane assembly 502. In other words, each other vane assembly 502 can be physically coupled to a retainer assembly 506 of the set of retainer assemblies 506. This configuration can reduce the total number of retainer assemblies 506 needed to effectively couple the disc 504 to the set of vane assemblies 502. Alternatively, the retainer assemblies 506 can be distributed across any number of vane assemblies 502. As a non-limiting example, the set of retainer assemblies 506 can be disposed along a portion of the disc 504 corresponding to each of a third, fourth, fifth, or nth vane assembly 502. Alternatively, the set of retainer assemblies 506 can be disposed along the disc 504 corresponding to each vane assembly 502 of the set of vane assemblies 502.

[0075] Benefits of the present disclosure include rotating blade assemblies that can be used with turboengines (e.g., counter-rotating turboengines) that have improved overall efficiency when compared to traditional turboengines (e.g., non-counter-rotating turboengines). For example, a traditional turboengine can include a set of rotating blades disposed downstream of a set of stationary vanes that collectively form a stage of a non-counter-rotating turboengine. During operation of the non-counter-rotating turboengine, a working airflow similar to that described herein can flow through the set of stationary vanes and subsequently flow to an adjacent set of rotating blades. The set of stationary vanes can be used to direct the working airflow to be incident with leading edges of the set of rotating blades, thereby limiting windage losses associated with non-incident working airflows. However, in the non-counter-rotating turboengine, work is extracted only by rotation of the set of rotating blades (e.g., the set of stationary vanes does not extract work from the working airflow). However, the present disclosure relates to a rotating blade assembly for a counter-rotating turboengine, where a stage is comprised of two adjacent rotating blade assemblies. It is contemplated that the adjacent rotating blade assemblies can rotate in opposite (e.g., opposing) circumferential directions relative to one another, however, the rotating blade assemblies can be positioned such that the working airflow exiting the upstream rotating blade assembly can be incident with respect to their own leading edges. As such, windage losses can still be avoided or otherwise limited, however, work can be extracted from both sets of rotating blade assemblies. This ultimately means that the total work output of the counter-rotating turboengine can be greater when compared to a similar sized non-counter-rotating turboengine (e.g., a similar or same number of total stages).

[0076] Further benefits of the present disclosure include rotating blade assemblies that have reduced losses when compared to rotating blade assemblies used within counter-rotating turboengines that do not have a disk as described herein. For example, a rotating blade assembly without a disk as described herein will form a non-rigid circumferential ring of blade assemblies within the working airflow. This in turn means that there is a greater tolerance of how much the blade assemblies can move during the contemplated circumferential movement (e.g., rotation) of the rotating blade assembly. The greater tolerance in turn results in the inner portion and outer band of the rotating blade assembly will be non-concentric, which ultimately creates losses. However, a rotating blade assembly as described herein includes a circumferential disk that interconnects each blade assembly within the rotating blade assembly. In other words, the disk can be used to form a rigid structure between adjacent blade assemblies. This in turn reduces or eliminates the gaps between adjacent components, which ensures concentricity between the disk and the outer band as described herein. The concentric rotation and assembly of the rotating blade assembly relative to the outer band in turn minimizes the losses created, which ultimately improves the efficiency of the rotating turboengine when compared to traditional rotating turboengines that do not have a rotating blade assembly as described herein.

[0077] Further benefits of the present disclosure include the rotating blade assembly within the counter-rotating turbine engine having increased frictional damping capability when compared to conventional rotating blade assemblies used within counter-rotating turbine engines. For example, conventional rotating blade assemblies can use a monolithic structure that interconnects adjacent blade assemblies. In other words, conventional rotating blade assemblies can include a single inner band formed as a single monolithic piece that is integral with the rest of the rotating blade assembly. However, the monolithic structure has low frictional damping capability because there are no contact surfaces between adjacent components. As a result, conventional rotating blade assemblies will vibrate, which in turn increases the overall losses associated with the operation of conventional counter-rotating turbine engines. However, the counter-rotating turbine engine as described herein includes a non-monolithic rotating blade assembly. As a non-limiting example, the counter-rotating turbine engine as described herein can include a non-monolithic disk that is coupled to and contacts the rest of the rotating blade assembly along various contact areas created by the interface between the rest of the rotating blade assembly and the disk (e.g., fastener assemblies, ribs, and tail / socket). These contact areas can create frictional contact areas between two adjacent portions in the rotating blade assembly. This in turn enhances the frictional damping capability of the rotating blade assembly when compared to conventional rotating blade assemblies. This reduces the vibrational losses associated with the operation of the counter-rotating turbine engine when compared to conventional counter-rotating turbine engines, which ultimately improves the overall efficiency of the counter-rotating turbine engine when compared to conventional counter-rotating turbine engines.

[0078] In areas not yet described, different features and structures of the various aspects can be used in combination with each other as desired. The fact that one feature is not described in all aspects does not mean that it cannot be interpreted, but for the sake of brevity of description. Therefore, various features of different aspects can be mixed and matched as desired to form new aspects, whether or not the new aspects are explicitly described. Combinations or permutations of features described herein are covered by the present disclosure.

[0079] This written description uses examples to describe the aspects of the disclosure described herein, including the best mode, and also enables one of ordinary skill in the art to practice the aspects of the disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the present disclosure is defined by the claims, and can include other examples that occur to those skilled in the art. Such other examples are intended to fall within the scope of the claims if they have structural elements in common with the

[0080] Further aspects of the present disclosure are provided by the subject matter of the following clauses.

[0081] A rotating blade assembly for a turbine engine having a drive shaft, the rotating blade assembly comprising: a disk operably coupled to the drive shaft and including a seat having at least a portion of a first through-hole; at least one blade assembly having an upper platform, a lower platform, a dovetail extending from the lower platform, and a blade extending between the upper platform and the lower platform; and a retainer assembly securing the disk to the blade assembly, the retainer assembly including: a hollow tubular element defining a second through-hole aligned with the first through-hole; and a pin extending through at least a portion of the first through-hole and the second through-hole and facing at least a portion of the hollow tubular element.

[0082] The rotating blade assembly of any of the preceding clauses, wherein the hollow tubular element is an inflatable bushing.

[0083] The rotating blade assembly of any of the preceding clauses, wherein the pin terminates at a distal end, and the rotating blade further includes a fastener secured to the distal end and abutting the disk.

[0084] The rotating blade assembly of any of the preceding clauses, wherein the fastener is one of a nut, a hydraulic fastener, a magnetic fastener, a weld, an adhesive, or an electrical connection.

[0085] The rotating blade assembly of any of the preceding clauses, wherein the fastener is a nut, and the pin includes a threaded section, wherein the nut is threadably connected to the threaded section of the pin.

[0086] The rotating blade assembly of any of the preceding clauses, wherein the dovetail defines at least another portion of the first through-hole.

[0087] The rotating blade assembly of any of the preceding clauses, wherein the disk further includes a first band and a second band collectively defining the seat.

[0088] The rotating blade assembly of any of the preceding clauses, wherein the first band and the second band are coupled at a lap joint.

[0089] The rotating blade assembly of any of the preceding clauses, wherein the first band includes a first rib and the second band includes a second rib, wherein the first rib and the second rib both face corresponding portions of the dovetail.

[0090] The rotating blade assembly of any of the preceding clauses, wherein the retainer assembly further includes a retainer plate including a through-hole aligned with the first through-hole and abutting the disk and a portion of the dovetail.

[0091] The rotating blade assembly of any of the preceding clauses, wherein the hollow tubular element further includes a first end including a shoulder, wherein the shoulder abuts a corresponding portion of the retainer plate.

[0092] The rotating blade assembly of any of the preceding clauses, wherein the hollow tubular element further comprises a shoulder, the shoulder abutting at least one of the disk or the dovetail.

[0093] The rotating blade assembly of any of the preceding clauses, wherein the disk comprises a dovetail connection, the dovetail connection extending from a remainder of the disk in a spanwise direction and extending through a corresponding portion of the dovetail.

[0094] The rotating blade assembly of any of the preceding clauses, wherein the disk extends continuously 360 degrees around the drive shaft.

[0095] A gas turbine engine comprising: an engine core defining an engine centerline and comprising a drive shaft and a first rotor; and a rotating blade assembly comprising: a disk operably coupled to the drive shaft and comprising a seat having at least a portion of a first through-hole; at least one blade assembly having an upper platform operably coupled to the first rotor, a lower platform, a dovetail extending from the lower platform, and a blade extending between the upper platform and the lower platform; and at least one retainer assembly securing the disk to the at least one blade assembly, the at least one retainer assembly comprising: a hollow tubular element defining a second through-hole, the second through-hole aligned with the first through-hole; and a pin extending through at least a portion of the first through-hole and the second through-hole and facing at least a portion of the hollow tubular element.

[0096] The gas turbine engine of any of the preceding clauses, wherein the at least one blade assembly is included within a set of blade assemblies, the set of blade assemblies being circumferentially spaced relative to one another and extending around an entirety of a circumferential direction of the engine centerline, and wherein the disk is a 360 degree ring extending circumferentially around each dovetail of the set of blade assemblies.

[0097] The gas turbine engine of any of the preceding clauses, wherein the at least one retainer assembly is included within a set of retainer assemblies, and wherein the set of retainer assemblies are disposed along the disk at circumferential locations corresponding to each other blade assembly, and wherein at least two adjacent blade assemblies define a socket extending at least partially through each dovetail, and wherein the disk further comprises a tail extending from a remainder of the disk and at least partially housed within the socket.

[0098] The gas turbine engine of any of the preceding clauses, further comprising a second rotor spaced radially outward from the first rotor relative to the engine centerline.

[0099] The gas turbine engine of any of the preceding clauses, wherein the disk further comprises a first band comprising a first rib facing the dovetail, the first band defining a first portion of the seat; and a second band comprising a second rib facing the dovetail, the second band defining a second portion of the seat, wherein the first rib and the second rib radially retain the disk to the dovetail.

[0100] The gas turbine engine of any of the preceding clauses, wherein the hollow tubular element is an inflatable bushing, the inflatable bushing further comprising a shoulder abutting at least one of the disk or the dovetail, and wherein the pin terminates at a distal end, the at least one retainer assembly further comprising a retainer plate comprising a through-hole aligned with the first through-hole and abutting a portion of the disk and the dovetail; and at least one fastener, the at least one fastener secured to the distal end and abutting the disk.

Claims

1. A rotating blade assembly rotatable about a rotation axis, characterized in that, The rotating blade assembly includes: A disk, the disk including a seat having at least a portion of a first through hole; A blade assembly having an upper platform, a lower platform, a dovetail extending from the lower platform, and a blade extending between the upper platform and the lower platform; and A retainer assembly that secures the disk to the blade assembly, the retainer assembly comprising: An expandable hollow tubular member defining a second through-hole aligned with a first through-hole, the expandable hollow tubular member comprising a tubular wall and a plurality of fingers extending axially relative to a rotation axis; and A pin that extends through at least a portion of the first through hole and the second through hole and faces at least a portion of the hollow tubular member.

2. The rotating blade assembly according to claim 1, characterized in that, in, The expandable hollow tubular component is an expandable bushing.

3. The rotating blade assembly according to claim 1, characterized in that, in, The pin terminates at the distal end, and the rotating blade assembly further includes fasteners fixed to the distal end and adjacent to the disk.

4. The rotating blade assembly of claim 1, wherein the expandable hollow tubular member comprises an expandable material.

5. The rotating blade assembly of claim 1, wherein the expandable hollow tubular member contacts a corresponding portion of the dovetail tenon.

6. The rotating blade assembly according to claim 1, characterized in that, in, The disc further includes a first band and a second band that jointly define the seat.

7. The rotating blade assembly according to claim 6, characterized in that, in, The first belt and the second belt are connected at the lap joint.

8. The rotating blade assembly according to claim 6, characterized in that, in, The first band includes a first rib, and the second band includes a second rib, wherein both the first rib and the second rib face the corresponding portion of the dovetail tenon.

9. The rotating blade assembly according to claim 1, characterized in that, in, The retainer assembly further includes a retainer plate having a through hole aligned with the first through hole and abutting a portion of the disc and the dovetail.

10. The rotating blade assembly according to claim 1, characterized in that, in, The expandable hollow tubular member further includes a shoulder that abuts at least one of the disc or the dovetail joint.

11. The rotating blade assembly according to claim 1, characterized in that, in, The disc includes a dovetail joint that extends from the rest of the disc in the span direction and through the corresponding portion of the dovetail.

12. A gas turbine engine, characterized in that, include: An engine core, which defines an engine centerline and includes a drive shaft and a first rotor; and The rotating blade assembly according to claim 1, wherein the disk is operatively coupled to the drive.

13. The gas turbine engine according to claim 12, characterized in that, in, The blade assembly is included within a plurality of blade assemblies, the set of blade assemblies being circumferentially spaced relative to each other, and wherein the disk is a 360-degree ring extending circumferentially around a corresponding portion of each of the plurality of blade assemblies.

14. The gas turbine engine according to claim 12, characterized in that, in, The blade assembly is included within a plurality of blade assemblies; The retainer assembly is included within a plurality of retainer assemblies, wherein the plurality of retainer assemblies secure the disk to the plurality of blade assemblies, and the plurality of retainer assemblies are disposed along the disk at circumferential positions corresponding to each of the plurality of blade assemblies. At least two adjacent blade assemblies of the plurality of blade assemblies include corresponding dovetail tenons, at least a portion of which is circumferentially spaced to define a mortise. The disc includes a tail portion that is at least partially housed within the socket.

15. The gas turbine engine according to claim 12, characterized in that, in, It further includes a second rotor that is radially spaced outward from the first rotor relative to the engine centerline.

16. A rotating blade assembly rotatable about a rotation axis, characterized in that, The rotating blade assembly includes: A disk, the disk including a seat having at least a portion of a first through hole; A blade assembly having an upper platform, a lower platform, a dovetail extending from the lower platform, and a blade extending between the upper platform and the lower platform; and A retainer assembly that secures the disk to the blade assembly, the retainer assembly comprising: An expandable hollow tubular member defining a second through-hole aligned with a first through-hole, the expandable hollow tubular member including a tubular wall defining the second through-hole, the tubular wall having a first section with an increased cross-sectional area, thereby defining a portion of the second through-hole with a reduced cross-sectional area when viewed along a plane extending along the axis of rotation and intersecting the expandable hollow tubular member; and A pin, which extends through at least a portion of the first through hole and the second through hole and faces at least a portion of the hollow tubular member, has a second section whose cross-sectional area gradually decreases when viewed along a plane that extends along the axis of rotation and intersects the pin.

17. The rotating blade assembly according to claim 16, characterized in that, in, The pin is a conical pin, and the expandable hollow tubular component is a conical hollow tubular component. The conical pin and the conical hollow tubular component form an interface.

18. The rotating blade assembly according to claim 16, characterized in that, in, The first segment and the second segment form an interface.

19. The gas turbine engine according to claim 17, characterized in that, The expandable hollow tubular component is an expandable bushing.

20. A rotating blade assembly for a turbine engine having a drive shaft, characterized in that, The rotating blade assembly includes: A disc, operably coupled to the drive shaft and including a seat having at least a portion of a first through hole, the disc including a dovetail joint extending from the remainder of the disc in the span direction; A blade assembly having an upper platform, a lower platform, a dovetail extending from the lower platform, and a blade extending between the upper and lower platforms, the dovetail connecting corresponding portions extending through the dovetail; and A retainer assembly that secures the disk to the blade assembly, the retainer assembly comprising: A hollow tubular member defining a second through-hole, the second through-hole being aligned with a first through-hole; and A pin that extends through at least a portion of the first through hole and the second through hole and faces at least a portion of the hollow tubular member.

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