阅读说明：本技术 涡轮发动机组件 (Turbine engine assembly ) 是由 皮奥特·耶日·库林斯基 托马斯·爱德华·伯多斯基 于 2020-09-04 设计创作，主要内容包括：用于涡轮发动机的组件可以包括整体式主体,该整体式主体具有内带和与内带径向地间隔开的外带,在内带或外带中的一个带中限定凹槽的袋状表面,以及具有从内带或外带中的另一个带径向地延伸的第一端的翼型件。(An assembly for a turbine engine may include a monolithic body having an inner band and an outer band radially spaced from the inner band, a pocket surface defining a groove in one of the inner or outer bands, and an airfoil having a first end extending radially from the other of the inner or outer bands.)

1. An assembly for a turbine engine, comprising:

a monolithic body comprising:

an inner band and an outer band radially spaced from the inner band;

a pocket surface defining a groove in one of the inner or outer bands; and

an airfoil including an outer wall defining a pressure side and a suction side, a first end extending radially from the other of the inner or outer bands, and a second end extending radially into the groove to define a gap between the airfoil and the one of the inner or outer bands.

2. The assembly of claim 1, further comprising a support structure at least partially closing the gap and coupled to at least one of the second end or the pocket surface.

3. The assembly of claim 2, wherein the support structure comprises:

an airfoil recess in the second end; and

a pin extending to the one of the inner band or the outer band and received in the airfoil groove to at least partially close the gap.

4. The assembly of claim 2, wherein the support structure comprises a clip spanning the gap and having a first clip end retained by the second end of the airfoil.

5. The assembly of claim 4, wherein the clip further comprises a spring clip having a second clip end retained by the one of the inner band or the outer band.

6. The assembly of claim 5, wherein the second end extends completely through the one of the inner band or the outer band.

7. The assembly of claim 4, wherein the clip further comprises a T-shaped body having a first leg terminating at the first clip end and extending to the second end of the airfoil.

8. The assembly of claim 7, wherein the T-shaped body includes a second leg extending perpendicular to the first leg and coupled to the one of the inner band or the outer band.

9. The assembly of claim 2, wherein the support structure comprises at least one curve damper connecting the second end of the airfoil with the one of the inner band or the outer band.

10. The assembly of any of claims 1-9, wherein the monolithic body further comprises vanes extending completely between the inner band and the outer band.

Technical Field

The present disclosure relates generally to airfoils within turbine engines that include static buckets or rotating blades, and more particularly, to turbine engine assemblies that include such airfoils.

Background

Turbine engines, particularly gas turbine engines or combustion turbine engines, are rotary engines that extract energy from a pressurized flow of combustion gases through the engine onto rotating turbine blades.

Gas turbine engines utilize mainstream flow to drive rotating turbine blades to produce thrust. The main flow is propelled by gas combustion to increase the thrust produced by the engine. Sealing members or other structures may be used to direct the airflow to desired locations within the engine, as well as to provide controlled movement of turbine engine components during operation.

Disclosure of Invention

In one aspect, the present disclosure is directed to an assembly for a turbine engine. The assembly includes a monolithic body having an inner band and an outer band radially spaced from the inner band, a pocket surface defining a groove in one of the inner or outer bands, and an airfoil including an outer wall defining a pressure side and a suction side, a first end extending radially from the other of the inner or outer bands, and a second end extending radially into the groove to define a gap between the airfoil and the one of the inner or outer bands.

In another aspect, the present disclosure is directed to a turbine engine. The turbine engine includes a compressor, a combustor, and an axially disposed turbine, and an assembly in at least one of the compressor or the turbine. The assembly includes a monolithic body having an inner band and an outer band radially spaced from the inner band, a pocket surface defining a groove in one of the inner or outer bands, and an airfoil including an outer wall defining a pressure side and a suction side, a first end extending radially from the other of the inner or outer bands, and a second end extending radially into the groove to define a gap between the airfoil and the one of the inner or outer bands.

In yet another aspect, the present disclosure is directed to an assembly for a turbine engine. The assembly includes a monolithic body having an inner band, an outer band radially spaced from the inner band, a groove in one of the inner or outer bands, and a second vane having a root at the other of the inner or outer bands and a second tip located within the groove and unaffected by the one of the inner or outer bands.

Drawings

In the drawings:

FIG. 1 is a schematic cross-sectional view of a turbine engine for an aircraft.

FIG. 2 is a perspective view of an assembly for the turbine engine of FIG. 1, according to various aspects described herein.

Fig. 3 is a side cross-sectional view of a portion of the assembly of fig. 2.

FIG. 4 is a side cross-sectional view of another assembly having a support structure that may be used in the turbine engine of FIG. 1, in accordance with various aspects described herein.

FIG. 5 is a side cross-sectional view of another assembly having another support structure that may be used in the turbine engine of FIG. 1, in accordance with various aspects described herein.

FIG. 6 is a side cross-sectional view of another assembly having another support structure that may be used in the turbine engine of FIG. 1, in accordance with various aspects described herein.

FIG. 7 is a side cross-sectional view of another assembly having another support structure that may be used in the turbine engine of FIG. 1, in accordance with various aspects described herein.

FIG. 8 is a side cross-sectional view of another assembly having another support structure that may be used in the turbine engine of FIG. 1, in accordance with various aspects described herein.

FIG. 9 is a side cross-sectional view of another assembly having another support structure that may be used in the turbine engine of FIG. 1, in accordance with various aspects described herein.

FIG. 10 is a side cross-sectional view of another assembly having another support structure that may be used in the turbine engine of FIG. 1, in accordance with various aspects described herein.

FIG. 11 is a side cross-sectional view of another assembly having another support structure that may be used in the turbine engine of FIG. 1, in accordance with various aspects described herein.

FIG. 12 is a side cross-sectional view of another assembly having another support structure that may be used in the turbine engine of FIG. 1, in accordance with various aspects described herein.

Detailed Description

The described embodiments of the present disclosure relate to an assembly for a turbine engine. For illustrative purposes, the present disclosure will be described with respect to an airfoil assembly of an aircraft turbine engine. However, it will be understood that the present disclosure is not so limited, and may have general applicability within any engine, as well as in non-aircraft applications (e.g., other mobile applications and non-mobile industrial, commercial, and residential applications).

Airfoil assemblies within turbine engines may be subjected to internal stresses during operation for a variety of reasons, including air pressure differences around each airfoil, thermal expansion of the airfoil or strip, or vibrational forces acting on the airfoil. Such stresses may occur anywhere along the airfoil, including at the attachment points or joints between the airfoil and the platform, band, or disk. In examples where stationary vanes are coupled at each end between the inner and outer bands, such internal stresses may cause component wear over time, such as fatigue, creep, or cracking of the airfoils.

As used herein, the term "forward" or "upstream" refers to movement in a direction toward the engine inlet, or toward a component that is relatively close to the engine inlet as compared to another component. The term "aft" or "downstream" as used in connection with "forward" or "upstream" refers to a direction toward the rear or outlet of the engine or a direction relatively closer to the engine outlet than another component.

As used herein, a "set" may include any number of the separately described elements, including only one element. Additionally, as used herein, the term "radial" or "radially" refers to a dimension extending between a central longitudinal axis of the engine and an outer circumference of the engine.

All directional references (e.g., radial, axial, proximal, distal, upper, lower, upward, downward, left, right, lateral, front, rear, top, bottom, above, below, vertical, horizontal, clockwise, counterclockwise, upstream, downstream, forward, rearward, etc.) are only used for identification purposes to aid the reader's understanding of the present disclosure, and do not constitute limitations, particularly as to the position, orientation, or use of the disclosure. Unless otherwise indicated, connection references (e.g., attached, coupled, connected, and engaged) are to be construed broadly and may include intermediate members between a collection of elements and relative movement between elements. As such, joinder references do not necessarily infer that two elements are directly connected and in fixed relation to each other. The exemplary drawings are for illustrative purposes only and the dimensions, locations, order and relative sizes reflected in the accompanying drawings may vary.

FIG. 1 is a schematic cross-sectional view of a gas turbine engine 10 for an aircraft. The engine 10 has a generally longitudinally extending axis or centerline 12, the axis or centerline 12 extending from a forward 14 to an aft 16. The engine 10 includes in downstream series flow relationship: a fan section 18 containing a fan 20, a compressor section 22 including a booster or Low Pressure (LP) compressor 24 and a High Pressure (HP) compressor 26, a combustion section 28 including a combustor 30, a turbine section 32 including a HP turbine 34 and a LP turbine 36, and an exhaust section 38.

The fan section 18 includes a fan housing 40 surrounding the fan 20. The fan 20 includes a plurality of fan blades 42 disposed radially about the centerline 12. The HP compressor 26, combustor 30, and HP turbine 34 form a core 44 of the engine 10, and the core 44 generates combustion gases. The core 44 is surrounded by a core housing 46, which core housing 46 may be coupled with the fan housing 40.

An HP shaft or spool 48, disposed coaxially about the centerline 12 of the engine 10, drivingly connects the HP turbine 34 with the HP compressor 26. An LP shaft or spool 50, disposed coaxially about the centerline 12 of the engine 10 within the larger diameter annular HP spool 48, drivingly connects the LP turbine 36 to the LP compressor 24 and fan 20. The spools 48, 50 are rotatable about the engine centerline and are coupled to a plurality of rotatable elements that may collectively define a rotor 51.

The LP and HP compressors 24, 26 each include a plurality of compressor stages 52, 54, with a set of compressor blades 56, 58 rotating relative to a corresponding set of stationary compressor vanes 60, 62 to compress or pressurize a fluid flow through the stages. In a single compressor stage 52, 54, a plurality of compressor blades 56, 58 may be arranged in an annular shape and may extend radially outward from the blade platform to the blade tip relative to the centerline 12, while corresponding static compressor vanes 60, 62 are positioned upstream of the rotating blades 56, 58 and adjacent to the rotating blades 56, 58. Note that the number of blades, vanes, and compressor stages shown in FIG. 1 is chosen for illustration purposes only, and other numbers are possible.

The vanes 56, 58 for the compressor stages may be mounted to (or integrated into) a disc 61, the disc 61 being mounted to a corresponding one of the HP spool 48 and the LP spool 50. The buckets 60, 62 for the compressor stages may be mounted to the core casing 46 in a circumferential arrangement.

The HP and LP turbines 34, 36 each include a plurality of turbine stages 64, 66, wherein a set of turbine blades 68, 70 rotate relative to a corresponding set of stationary turbine buckets 72, 74 (also referred to as nozzles) to extract energy from the fluid flow through the stages. In a single turbine stage 64, 66, a plurality of turbine blades 68, 70 may be arranged in a ring shape and may extend radially outward relative to the centerline 12, while corresponding static turbine vanes 72, 74 are positioned upstream of and adjacent to the rotating blades 68, 70. Note that the number of blades, buckets, and turbine stages shown in FIG. 1 is chosen for illustrative purposes only, and other numbers are possible.

The blades 68, 70 for the turbine stages may be mounted to a disc 71, the disc 71 being mounted to a corresponding one of the HP spool 48 and the LP spool 50. The buckets 72, 74 for the compressor stages may be mounted to the core housing 46 in a circumferential arrangement.

In addition to the rotor portion, stationary portions (stationality) of the engine 10, such as the static vanes 60, 62, 72, 74 in the compressor section 22 and the turbine section 32, are also referred to individually or collectively as the stator 63. As such, the stator 63 may refer to a combination of non-rotating elements throughout the engine 10.

In operation, the airflow exiting fan section 18 is divided such that a portion of the airflow is channeled into LP compressor 24, which LP compressor 24 subsequently supplies pressurized air 76 to HP compressor 26, which HP compressor 26 further pressurizes the air. Pressurized air 76 from the HP compressor 26 is mixed with fuel in the combustor 30 and ignited, thereby generating combustion gases. The HP turbine 34 extracts some work from these gases, driving the HP compressor 26. The combustion gases are discharged into the LP turbine 36, the LP turbine 36 extracts additional work to drive the LP compressor 24, and the exhaust gases are ultimately discharged from the engine 10 via an exhaust section 38. The LP turbine 36 is driven to drive the LP spool 50 to rotate the fan 20 and LP compressor 24.

A portion of the pressurized airflow 76 may be withdrawn from the compressor section 22 as bleed air (bleedair) 77. This bleed air 77 may be withdrawn from the pressurized airflow 76 and provided to engine components that require cooling. The temperature of the pressurized gas stream 76 entering the combustor 30 increases significantly. As such, the cooling provided by the bleed air 77 is necessary to operate such engine components in an elevated temperature environment.

The remaining portion of airflow 78 bypasses LP compressor 24 and engine core 44 and exits engine assembly 10 on a fan discharge side 84 through a row of static vanes, more specifically an outlet guide vane assembly 80, which outlet guide vane assembly 80 includes a plurality of airfoil guide vanes 82. More specifically, a circumferential row of radially extending airfoil guide vanes 82 is utilized adjacent fan section 18 for some directional control of airflow 78.

Some of the air supplied by fan 20 may bypass engine core 44 and be used to cool portions of engine 10, particularly hot portions, and/or to cool or power other aspects of the aircraft. In the case of a turbine engine, the hot portion of the engine is generally downstream of the combustor 30, particularly the turbine section 32, with the HP turbine 34 being the hottest portion, as the HP turbine 34 is directly downstream of the combustion section 28. Other sources of cooling fluid may be, but are not limited to, fluid discharged from the LP compressor 24 or the HP compressor 26.

Referring now to FIG. 2, an assembly 100 that may be used in the turbine engine 10 of FIG. 1 is shown. The assembly 100 is shown in the form of an airfoil assembly. For example, the assembly 100 may include the HP turbine bucket 72 (FIG. 1). It will be appreciated that aspects of the present disclosure may also be applicable to rotating blades or static buckets, and that the assembly 100 may be located at any suitable location within the turbine engine 10, including but not limited to the compressor section 22 or the turbine section 32.

The assembly 100 may include a body 101. It is contemplated that the body 101 of the assembly 100 may include a plurality of separate components, such as rotating blades, static vanes, platforms, inner or outer bands, etc., assembled together, for example, using attachment hardware or other joining methods. In the example of fig. 2, the body 101 is in the form of a unitary body 101. As used herein, "unitary body" shall mean a body that is formed as a single, unitary piece. Such a monolithic body may comprise multiple parts or elements, and it will be understood that such parts or elements may also be formed from the monolithic body without further attachment via hardware, adhesive, or the like. For example, a body formed from a plurality of individual elements secured together (e.g., adhesive, bolts, etc.) does not form a unitary body.

The monolithic body 101 includes an inner band 102 and an outer band 103 radially spaced from the inner band 102. The inner and outer bands 102, 103 may at least partially define a circumferential direction C, a radial direction R, and an axial direction a, as shown. The axial direction a may be aligned with the engine centerline 12 (fig. 1).

Pocket surface 104 defining groove 105 may be included in one or both of inner band 102 or outer band 103. In the example shown, pocket surface 104 and groove 105 are located in outer band 103. As used herein, a "baggy surface" will refer to a portion of a wall that is curved or bent to define a groove within the wall.

The monolithic body 101 may also include vanes 110 that extend completely between the inner band 102 and the outer band 103. The bucket 110 includes a bucket outer wall 111, the bucket outer wall 111 having a bucket leading edge 112 and a bucket trailing edge 113 and defining a bucket pressure side 114 and a bucket suction side 115. The bucket outer wall 111 may also define radially spaced first and second ends, shown as bucket root 116 at the inner band 102 and bucket tip 117 at the outer band 103, respectively.

The monolithic body 101 may also include an airfoil 120, the airfoil 120 having an outer wall 121, the outer wall 121 having a leading edge 122 and a trailing edge 123 and defining a pressure side 124 and a suction side 125. The outer wall 121 may also define radially spaced first and second ends, shown as root 126 and tip 127, respectively.

In the illustrated example, the airfoil 120 is in the form of a static vane. It will be understood that aspects of the present disclosure may also be applied to a rotating blade extending from a platform at the root, where either the inner band or the outer band may form the platform. A radial length 128R of airfoil 120 is defined between second root 126 and second tip 127. An axial length 128A of airfoil 120 is defined between second leading edge 122 and second trailing edge 123, as shown.

Furthermore, the tip 127 extends radially into the groove 105. A gap 106 may be defined between the tip 127 and the pocket surface 104. In the example shown, the second root 126 extends from the inner band 102 and the tip 127 extends into the groove 105 in the outer band 103.

An exemplary heated airflow 130 is shown, the heated airflow 130 moving through the assembly 100, through the bucket 110 and the airfoil 120. An exemplary force F is shown in a circumferential direction on airfoil 120 due to heating airflow 130. The curvature of pressure side 124 and suction side 125 (FIG. 2) results in a lower air pressure near suction side 125 than pressure side 124, resulting in a total force F acting on airfoil 120, as shown. This force F is also referred to as the "lift" force on the airfoil. It will be appreciated that, although not shown, the force F from the heated airflow 130 may also be applied to the bucket 110 in a manner similar to that described for the airfoil 120.

It will be appreciated that the assembly 100 may be formed as a single piece having a unitary body 101 with inner and outer bands 102, 103, pocket surfaces 104, vanes 110, and airfoils 120. Any suitable manufacturing method or process, including casting or additive manufacturing, may be used to form the component 100. As used herein, an "additive manufactured" part will refer to a part formed by an Additive Manufacturing (AM) process, wherein the part is built layer by continuously depositing materials. AM is a proper name to describe a technique for building 3D objects from layer-by-layer added material (whether the material is plastic or metal). AM technology can utilize computers, 3D modeling software (computer aided design or CAD), machine equipment and layered materials. Once the CAD sketch is generated, the AM equipment can read the data from the CAD file and place or add successive layers of liquid, powder, sheet material, or other material layer by layer to make the 3D object. It should be understood that the term "additive manufacturing" encompasses many technologies, including such subsets as 3D printing, Rapid Prototyping (RP), Direct Digital Manufacturing (DDM), layered manufacturing, and additive manufacturing. Non-limiting examples of additive manufacturing that may be used to form an additive manufactured part include powder bed melting, photocuring, adhesive jetting, material extrusion, directed energy deposition, material jetting, or sheet lamination.

In the example shown in FIG. 2, the assembly 100 includes a pair of airfoils 120 with buckets 110 spaced between the pair of airfoils 120. A corresponding pair of pocket surfaces 104 and grooves 105 are provided in the outer band 103, and the tip 127 of each airfoil 120 extends into each groove 105, as shown. It is contemplated that multiple assemblies 100 may be coupled together circumferentially to form an annular structure about engine centerline 12, such as a portion of HP turbine stage 64 (FIG. 1). In an alternative example, the unitary body of the assembly may include an entire annular structure about the centerline 12 having annular inner and outer bands with the buckets 110 and airfoils 120.

In another example, it is contemplated that the pocket-like surface and the groove may be located in the inner band. In this case, the root may extend radially from the outer band and the tip may extend into a groove located in the inner band. In yet another example (not shown), pocket-like surfaces and grooves may be provided on both the inner and outer bands. In this case, one non-limiting embodiment may include an alternative airfoil in which the bucket extends fully between the inner and outer bands along the circumferential direction, the airfoil having a root extending from the outer band and a tip extending into a groove in the inner band, a third airfoil extending fully between the inner and outer bands, and a fourth airfoil having a root extending from the inner band and a tip extending into a groove in the outer band.

In yet another example, the monolithic body may include a plurality of airfoils each extending from a root of one of the inner or outer bands to a tip extending into a groove in the other of the inner or outer bands and without additional vanes included in the assembly. In this example, each airfoil in the monolithic body includes a tip that extends into a corresponding groove. The grooves may be formed on the inner band only, the outer band only, or both the inner and outer bands.

Turning to fig. 3, a cross-sectional view of the assembly 100 is shown adjacent the groove 105. In the example shown, the pocket surface 104 forms a generally U-shaped groove 105, the groove 105 having a U-shaped gap 106 from a tip 127 of the airfoil 120. Gap 106 may define a gap distance 107 between any portion of pocket surface 104 and tip 127. For clarity, one exemplary gap distance 107 is shown at the radially outermost portion of the tip 127. In one example, the gap distance 107 may be 0.5-0.8 mm. Additionally, it is contemplated that gap distance 107 may vary along different portions of gap 106. In the example shown, a first clearance distance 107A is defined between tip 127 and pocket surface 104 adjacent pressure side 124, and a second clearance distance 107B is defined between tip 127 and pocket surface 104 adjacent suction side 125. The second gap distance 107B may be greater than, less than, or equal to the first gap distance 107A.

Optionally, a support structure 140 may be included in the assembly 100. The support structure 140 may at least partially close the gap 106, including completely closing the gap 106. The support structure 140 may be coupled to at least one of the tip 127 or the pocket surface 104. For example, the support structure 140 may be in contact with one of the pressure side 124 or the suction side 125 of the airfoil 120. Further, the support structure 140 may extend along at least a portion of the axial length 128A (fig. 2) of the airfoil 120, including the entire axial length 128A.

During operation, the heated airflow 130 and the elevated ambient temperature may cause the airfoil 120 to begin thermal expansion from its unheated or baseline state. Accordingly, one or both of the radial length 128R or the axial length 128A (fig. 2) may be increased and the gap distance 107 may be decreased. Additionally, vibration or other forces may cause movement of the tip 127 within the groove 105. During operation of the engine 10, the force F may also urge the tip 127 toward one side of the groove 105, thereby causing the tip 127 to form a gap distance 107B that is less than the first gap distance 107A. It will be appreciated that allowing the tip 127 to move freely within the groove 105 may reduce stresses within the airfoil 120 during operation as compared to holding both ends of a conventional static bucket between the inner and outer bands. In examples where the assembly 100 includes a support structure 140, such a support structure 140 may provide for controlled or partially constrained movement of the tip 127 while still allowing for free movement during operation.

Referring now to FIG. 4, a cross-sectional view of another assembly 200 that may be used in the turbine engine 10 of FIG. 1 is shown. Assembly 200 is similar to assembly 100. Accordingly, like components will be described with like numerals incremented by 100 and it will be understood that the description of like components of the assembly 100 applies to the assembly 200 unless otherwise noted.

The assembly 200 includes a body 201, the body 201 having an inner band (not shown) similar to the inner band 102, an outer band 203, buckets (not shown) similar to the buckets 110, and an airfoil 220 similar to the airfoil 120, and the airfoil 220 having a pressure side 224, a suction side 225, a first end (not shown) similar to the root 126, and a second end in the form of a tip 227. It is contemplated that the body 201 may be in the form of a unitary body 201 as described above. Outer band 203 includes a pocket surface 204 that defines a groove 205, and a gap 206 is defined between tip 227 and pocket surface 204. During operation of engine 10 (FIG. 1), as described above, total force F acts on airfoil 220 as a result of airflow through assembly 200.

A support structure 240 is included in the assembly 200. In the illustrated example, the support structure 240 includes an airfoil groove 242 in the tip 227, a strap aperture 244 in the outer strap 203, and a pin 246 extending into or through the strap aperture 244 and received in the airfoil groove 242. In one example, the airfoil groove 242 and the strap aperture 244 may be formed in the same manufacturing process as the monolithic body 201. In another example, the airfoil groove 242 and the band hole 244 may be formed by drilling, for example, a precision hole through the outer band 203 and into the tip 227. In this case, the pin 246 may be inserted into the hole 244. In one example, pin 246 may be retained within aperture 244 by an interference fit, wherein aperture 244 is slightly smaller than the diameter of pin 246 such that pin 246 is pressed into aperture 244 under pressure. Additionally or alternatively, pin 246 may be secured by spot welding, adhesive, a secondary locking pin (not shown) extending vertically through pin 246, or the like, or any combination thereof. The pin 246 may also at least partially close the gap 206. In the example of fig. 4, the pin 246 completely closes the gap 206 at the axial position of the pin 246. In the example shown, air cannot flow completely around the tip 227 through the groove 205 in the plane of the cross-sectional view, but such flow may occur upstream or downstream of the pin 246. Optionally, a radial space 248 may be defined between the pin 246 and the airfoil 220 such that the airfoil 220 may be free to radially expand, for example, by thermal expansion.

It is also contemplated that in examples utilizing drilling to form airfoil groove 242 or strap aperture 244, monolithic body 201 may include a securing ligament (not shown) to temporarily hold or support tip 227 during the drilling process. Additionally or alternatively, the gap 206 may be filled with a sacrificial or temporary material (e.g., hard wax) for support. Such retaining ligaments may be formed at the location of the airfoil 220 and the outer band 203, and then removed, and any sacrificial or temporary support material may be subsequently removed, such as by melting or dissolving.

Further, the airfoil groove 242, the belt hole 244, and the pin 246 are shown as having a substantially rectangular geometric profile, for example, to accommodate a cylindrical pin 246. In another example (not shown), the diameter of the pin or the hole may vary in the radial direction R. For example, the pin adjacent the tip of the airfoil may have a narrower diameter and a wider diameter within the outer band. For example, such pins having a conical or frustoconical geometric profile may be used. Any geometric profile is contemplated for use with airfoil groove 242, belt holes 244 and pin 246.

During operation of engine 10, force F may urge tip 227 circumferentially against pin 246 via airfoil groove 242, which may further act to retain pin 246. The airfoil 220 may still be provided with radial or circumferential freedom of movement, as shown, and the pin 246 may provide a degree of fluid sealing in the gap 206 between the pressure side 224 and the suction side 225.

Referring now to FIG. 5, a cross-sectional view of another assembly 300 that may be used in the turbine engine 10 of FIG. 1 is shown. The assembly 300 is similar to the assemblies 100, 200. Accordingly, like parts will be described with like numerals further incremented by 100, with the understanding that the description of like parts of the assemblies 100, 200 applies to the assembly 300 unless otherwise noted.

The assembly 300 includes a body 301 having an inner band (not shown) similar to the inner band 102, an outer band 303, buckets (not shown) similar to the buckets 110, and an airfoil 320, the airfoil 320 having a pressure side 324, a suction side 325, a first end (not shown) similar to the root 126, and a second end in the form of a tip 327. It is contemplated that the body 301 may be in the form of a unitary body 301, as described above. Outer band 303 includes a pocket surface 304 that defines a groove 305 and defines a gap 306 between a tip 327 and pocket surface 304. During operation of engine 10 (FIG. 1), as airflow passes through assembly 300, a total force F acts on airfoil 320, as described above.

A support structure 340 is included in the assembly 300. One difference compared to the assemblies 100, 200 is that the support structure 340 is in the form of a damper 350 connecting the tip 327 and the pocket surface 304. Damper 350 may have a curvilinear geometric profile such that damper 350 forms a spring-like or resilient structure. In this case, during operation of engine 10 (FIG. 1), circumferential or radial movement of tip 327 may be inhibited while still providing thermal expansion of airfoil 320. Additionally, damper 350 may completely close gap 306 such that air cannot flow through the groove between pressure side 324 and suction side 325. Alternatively, the damper 350 may be formed with perforations or other apertures (not shown), for example, for weight reduction while still providing an effective seal.

It is further contemplated that damper 350 may extend axially along tip 327 and pocket surface 304. In one example, damper 350 may extend along the entire axial length of airfoil 320, as described above with respect to fig. 2. In another example, a plurality of axially spaced dampers 350 may be disposed between the airfoil 320 and the pocket surface 304.

Turning to FIG. 6, a cross-sectional view of another assembly 400 that may be used in the turbine engine 10 of FIG. 1 is shown. The assembly 400 is similar to the assemblies 100, 200, 300. Accordingly, like parts will be described with like numerals further incremented by 100, with the understanding that the description of like parts of the assemblies 100, 200, 300 applies to the assembly 400 unless otherwise noted.

The assembly 400 includes a body 401 having an inner band (not shown) similar to the inner band 102, an outer band 403, a bucket (not shown) similar to the bucket 110, and an airfoil 420 having a pressure side 424, a suction side 425, a first end (not shown) similar to the root 126, and a second end in the form of a tip 427. It is contemplated that the body 401 may be in the form of a unitary body 401, as described above. The outer band 403 includes a pocket surface 404 defining a groove 405 and a gap 406 is defined between the tip 427 and the pocket surface 404. During operation of engine 10 (FIG. 1), a total force F acts on airfoil 420, as described above.

A support structure 440 is included in the assembly 400. One difference compared to the assemblies 100, 200, 300 is that the support structure 440 is in the form of a clip 452 having a first clip end 454 and a second clip end 456. The clip 452 may be formed of any suitable material, including but not limited to metal, polymer, or composite materials.

It is also contemplated that the clip 452 may extend axially along the tip 427 and the pocket surface 404. In one example, the clip 452 may extend along the entire axial length of the airfoil 420, as described above with respect to fig. 2. In another example, a plurality of axially spaced clips 452 may be provided between the airfoil 420 and the pocket surface 404.

Airfoil 420 includes an airfoil flute 442, the airfoil flute 442 forming a lip 443 on suction side 425. The clip 452 has a curvilinear geometric profile forming a spring clip 452 with a first clip end 454 retained by the tip 427 and a second clip end 456 retained by the outer band 403. More specifically, the first clip end 454 may be retained by the lip 443 and the second clip end 456 may be retained by the brace 458 coupled to the outer band 403. In one example, the stent 458 may be included in the outer band 403 as part of the monolithic body 401.

When assembled, as shown, clip 452 may expand circumferentially into place and across gap 406 such that air cannot flow through the groove between pressure side 424 and suction side 425. Alternatively, the clips 452 may be formed with perforations or other holes (not shown), for example, for weight reduction while still providing an effective seal. During operation of the engine 10 (FIG. 1), circumferential or radial movement of the tips 427 may be inhibited by the spring clips 452 while still providing thermal expansion of the airfoil 420.

FIG. 7 illustrates a cross-sectional view of another assembly 500 that may be used in the turbine engine 10 of FIG. 1. The assembly 500 is similar to the assemblies 100, 200, 300, 400. Accordingly, like parts will be described with like numerals further incremented by 100, it being understood that the description 300 of like parts of the assemblies 100, 200, 300, 400 applies to the assembly 500 unless otherwise noted.

The assembly 500 includes a body 501, the body 501 having an inner band (not shown) similar to the inner band 102, an outer band 503, a bucket (not shown) similar to the bucket 110, and an airfoil 520, the airfoil 520 having a pressure side 524, a suction side 525, a first end (not shown) similar to the root 126, and a second end in the form of a tip 527. It is contemplated that the body 501 may be in the form of a unitary body 501, as described above. Outer band 503 includes a pocket surface 504 that defines a groove 505, and defines a gap 506 between a tip 527 and pocket surface 504. During operation of engine 10 (FIG. 1), total force F acts on airfoil 520, as described above.

A support structure 540 is included in the assembly 500. The support structure 540 is in the form of a clip 552 having a first clip end 554 and a second clip end 556. The clip 552 may be formed from any suitable material, including but not limited to metal, polymer, or composite materials.

The clip 552 may also extend axially along the tip 527 and the pocket surface 504. In one example, the clip 552 may extend along the entire axial length of the airfoil 520, as described above with respect to fig. 2. In another example, a plurality of axially spaced clips 552 may be disposed between the airfoil 520 and the pocket surface 504.

One difference compared to assemblies 100, 200, 300, 400 is that outer band 503 includes a band aperture 544 through which a clip 552 may extend through aperture 544. In the example shown, the first clip end 554 has a curved profile and is retained between the tip 527 and the pocket surface 504. Clip 552 extends completely through outer band 503 via strap aperture 544 and second clip end 556 is coupled to outer band 503. For example, clips 552 may be inserted through the strap apertures 544 and retained via the first clip end 554, and the attachment points 560 may be used to secure the second clip end 556 to the outer strap 503, such as by spot welding, brazing, adhesives, locking pins, or other attachment mechanisms.

When assembled, the clip 552 may span the gap 506 such that air cannot flow through the groove between the pressure side 524 and the suction side 525. During operation of the engine 10 (fig. 1), circumferential or radial movement of the tip 527 may be controlled by the clip 552 while still providing for thermal expansion of the airfoil 420.

FIG. 8 illustrates a cross-sectional view of another assembly 600 that may be used in the turbine engine 10 of FIG. 1. The assembly 600 is similar to the assemblies 100, 200, 300, 400, 500. Accordingly, like parts will be described with like numerals further incremented by 100, it being understood that the description of like parts of the assemblies 100, 200, 300, 400, 500 applies to the assembly 600 unless otherwise noted.

The assembly 600 includes a body 601 having an inner band (not shown) similar to the inner band 102, an outer band 603, buckets (not shown) similar to the buckets 110, and an airfoil 620, the airfoil 620 having a pressure side 624, a suction side 625, a first end (not shown) similar to the root 126, and a second end in the form of a tip 627. It is contemplated that body 601 may be in the form of a unitary body 601, as described above. Outer band 603 includes a pocket surface 604 defining a groove 605, and a gap 606 is defined between tip 627 and pocket surface 604. During operation of engine 10 (FIG. 1), a total force F acts on airfoil 620, as described above.

A support structure 640 is included in the assembly 600. The support structure 640 is in the form of a clip 652. One difference compared to the assemblies 100, 200, 300, 400, 500 is that the clip 652 includes a T-shaped body 662, the T-shaped body 662 having a first leg 664 terminating at a first clip end 654. The first leg 664 extends through a hole 644 in the outer band 403, across the gap 606, and into the tip 627 via an airfoil groove 642. In addition, the T-shaped body 662 includes a second leg 666 that extends perpendicular to the first leg 664. Second leg 666 may be coupled to outer band 603. In the example shown, a securing pin 668 is used to secure the second leg 666. In other examples (not shown), second leg 666 may be welded, brazed, secured with an adhesive, or secured with other hardware, such as bolts, screws, locking pins, and the like. The clips 652 may also provide for radial or circumferential movement of the airfoil 620 during operation of the engine 10 (FIG. 1) as described above.

When assembled, support structure 640 may form a labyrinth seal 665 defined at least in part by gap 606 and first leg 664. As is generally understood by labyrinth seals, a small amount of air may be able to flow around the labyrinth path defined by labyrinth seal 665 while preventing a large amount of air from flowing past seal 665. For example, a small amount of air may flow around the first leg 664 and the first clip end 652 from the pressure side 624 into the groove 605, into the airfoil groove 642, and out the groove 605 at the suction side 625.

FIG. 9 illustrates a cross-sectional view of another assembly 700 that may be used in the turbine engine 10 of FIG. 1. The assembly 700 is similar to the assemblies 100, 200, 300, 400, 500, 600. Accordingly, like parts will be described with like numerals further incremented by 100, with the understanding that the description of like parts of assemblies 100, 200, 300, 400, 500, 600 applies to assembly 700 unless otherwise noted.

The assembly 700 includes a body 701 having an inner band (not shown) similar to the inner band 102, an outer band 703, a bucket (not shown) similar to the bucket 110, and an airfoil 720, the airfoil 720 having a pressure side 724, a suction side 725, a first end (not shown) similar to the root 126, and a second end in the form of a tip 727. It is contemplated that the body 701 may be in the form of a unitary body 701, as described above. The outer band 703 includes a pocket surface 704 that defines a groove 705, and a gap 706 is defined between the tip 727 and the pocket surface 704. During operation of engine 10 (FIG. 1), total force F acts on airfoil 720, as described above.

A support structure 740 is included in the assembly 700. The support structure 740 may be coupled by at least one coupling point 760. The support structure 740 may also extend axially along the tip 727 and pocket surface 704. In one example, support structure 740 may extend along the entire axial length of airfoil 720, as described above with respect to fig. 2. In another example, a plurality of axially spaced support structures 740 may be provided between the airfoil 720 and the pocket surface 704.

One difference compared to assemblies 100, 200, 300, 400, 500, 600 is that support structure 740 is configured to disengage at least one coupling point 760 during operation of engine 10 (fig. 1) or as a result of purposeful operation prior to assembly. In the example of fig. 9, two coupling points 760 are provided between the outer band 703 and the nib 727. The attachment point 760 may be in the form of a ligament or protrusion 761 formed with the unitary body 701, the unitary body 701 being formed with the outer band 703 and the airfoil 720. During operation, the coupling point 760 may disengage from the nib 727 to define a narrow gap 780. In non-limiting examples, such detachment may occur due to thermal expansion of airfoil 720, radial or circumferential movement of tip 727, or due to an applied load (not shown).

The narrow gap 780 may be smaller or narrower than the gap 706. In one example of forming monolithic body 701 by additive manufacturing, it is contemplated that gap 706 may be 500 microns between pocket surface 704 and second end 727. In this case, the narrow gap 780 may be less than 500 microns, including 5-50 microns in non-limiting examples. It is contemplated that the narrow gap 780 formed by the disengaged coupling points 760 may be smaller than currently available through additive manufacturing or other forming processes. Additionally, such a narrow gap 780 may provide a seal of gap 706, wherein only a very small amount of air may flow from pressure side 724 through groove 705 to suction side 725.

FIG. 10 illustrates a cross-sectional view of another assembly 800 that may be used in the turbine engine 10 of FIG. 1. The assembly 800 is similar to the assemblies 100, 200, 300, 400, 500, 600, 700. Accordingly, like parts will be described with like numerals further incremented by 100, it being understood that the description of like parts of the assemblies 100, 200, 300, 400, 500, 600, 700 applies to the assembly 800, unless otherwise noted.

The assembly 800 includes a body 801 having an inner band (not shown) similar to the inner band 102, an outer band 803, a bucket (not shown) similar to the bucket 110, and an airfoil 820, the airfoil 820 having a pressure side 824, a suction side 825, a first end (not shown) similar to the root 126, and a second end in the form of a tip 827. It is contemplated that body 801 may be in the form of a unitary body 801, as described above. The outer band 803 includes a pocket surface 804 that defines a groove 805, and a gap 806 is defined between the tip 827 and the pocket surface 804. During operation of engine 10 (FIG. 1), a total force F acts on airfoil 820, as described above.

A support structure 840 is included in the assembly 800. The support structure 840 may be coupled by at least one coupling point 860, the at least one coupling point 860 configured to disengage during operation of the engine 10 (fig. 1).

In the example of fig. 10, two attachment points 860 are provided between the outer band 803 and the tip 827. Attachment point 860 may be in the form of a ligament or protrusion 861 defined by pocket surface 804. One difference compared to assemblies 100, 200, 300, 400, 500, 600, 700 is that the protrusion 861 is located radially outward from the opening to the groove 805, as shown. During operation, the coupling point 860 may disengage from the tip 827 to define a narrow gap 880, which narrow gap 880 may be less than or narrower than the gap 806 as described above.

FIG. 11 illustrates a cross-sectional view of another assembly 900 that may be used in the turbine engine 10 of FIG. 1. The assembly 900 is similar to the assemblies 100, 200, 300, 400, 500, 600, 700, 800. Accordingly, like components will be described with like numerals further incremented by 100, with the understanding that the description of like components of the assemblies 100, 200, 300, 400, 500, 600, 700, 800 applies to the assembly 900 unless otherwise noted.

The assembly 900 includes a body 901, the body 901 having an inner band (not shown) similar to the inner band 102, an outer band 903, a bucket (not shown) similar to the bucket 110, and an airfoil 920, the airfoil 920 having a pressure side 924, a suction side (not shown) similar to the suction sides 125, 225, 325, 425, 525, 625, 725, 825, a first end (not shown) similar to the root 126, and a second end in the form of a tip 927. It is contemplated that the body 901 may be in the form of a unitary body 901, as described above. The outer band 903 includes a pocket-like surface 904 defining a recess 905, and a gap 906 is defined between the tip 927 and the pocket-like surface 904. During operation of engine 10 (FIG. 1), a total force F acts on airfoil 920, as described above.

A support structure 940 is included in the assembly 900. Support structure 940 may be coupled by at least one coupling point configured to disengage during operation of engine 10 (fig. 1). One difference compared to the assemblies 100, 200, 300, 400, 500, 600, 700, 800 is that the support structure 940 includes a central body 982, the central body 982 having a first coupling point 983 at the outer band 903 and a second coupling point 984 at the tip 927. Although the central body 982 is shown as having a circular profile, such as a spherical or cylindrical central body 982, any geometric profile is contemplated for the support structure 940. During operation, the coupling points 983, 984 may disengage from the respective outer band 903 and tip 827 to define a corresponding narrow gap 980, which narrow gap 980 may be less than or narrower than the gap 906 as described above. More specifically, the center body 982 may be carried by the airflow over the tip 927 from the pressure side 924 toward the suction side 925. Due to the airflow, the central body 982 may wedge into the gap 906 to seal the gap 906. The outer band surface 904 and the wedge-shaped surface of the airfoil tip 927 may be arranged as shown in fig. 11, but are not limited to the configuration or geometric profile shown. The surfaces around the groove 905 may form any wedge-shaped configuration with the central body 982 that will seal the flow of the airfoil 920 or inhibit movement of the airfoil 920 while allowing thermal expansion.

Turning to FIG. 12, a cross-sectional view of another assembly 1000 that may be used in the turbine engine 10 of FIG. 1 is shown. The assembly 1000 is similar to the assemblies 100, 200, 300, 400, 500, 600, 700, 800, 900. Accordingly, like components will be described with like numerals further incremented by 100, with the understanding that the description of like components of assemblies 100, 200, 300, 400, 500, 600, 700, 800, 900 applies to assembly 1000 unless otherwise noted.

The assembly 1000 includes a body 1001 having an inner band (not shown) similar to the inner band 102, an outer band 1003, a bucket (not shown) similar to the bucket 110, and an airfoil 1020, the airfoil 1020 having a pressure side 1024, a suction side 1025, a first end (not shown) similar to the root 126, and a second end in the form of a tip 1027. It is contemplated that body 1001 may be in the form of a unitary body 1001, as described above. The outer band 1003 includes a pocket surface 1004 that defines a groove 1005, and a gap 1006 is defined between the tip 1027 and the pocket surface 1004. During operation of engine 10 (FIG. 1), a total force F acts on airfoil 1020, as described above.

A support structure 1040 is included in the assembly 1000. The support structure 1040 is in the form of a protrusion 1061 defined by the pocket surface 1004 and in contact with the second end 1027 of the airfoil 1020. One difference compared to assemblies 100, 200, 300, 400, 500, 600, 700, 800, 900 is that protrusion 1061 may be formed via deformation of unitary body 1001, such as with tool 1090. In one example, the tool 1090 may be in the form of a circular punch that plastically deforms the outer band 1003 and forms the protrusion 1061 in contact with the second end 1027. Tool 1090 can include any type of punch, etc. to form protrusion 1061.

The protrusion 1061 may also extend axially along the tip 1027 and the pocket surface 1004. For example, the protrusion 1061 may extend along the entire axial length of the airfoil 1020, as described above with respect to fig. 2. In another example, a plurality of axially spaced protrusions 1061 may be formed in the pocket surface 1004. Although shown along the suction side 1025, the protrusions 1061 may be formed at any location along the groove 1005, including radially outward along the pressure side 1024 or the tip 1027. During operation of engine 10 (fig. 1), protrusions 1061 may provide sealing of gap 1006 and controlled circumferential or radial movement of tip 1027, as described above.

It will be appreciated that the above aspects of the assemblies 100, 200, 300, 400, 500, 600, 700, 800, 900, 1000 may be combined with or substituted for one another to form various other aspects. Some non-limiting examples will be described below, and the present disclosure still covers other examples of combinations or alternatives not explicitly described.

In one example, the assembly may include a curvilinear damper in addition to the point of decoupling between the airfoil and the pocket surface. In this case, the curved damper may provide controlled movement of the second end after sufficient force is applied to disengage the coupling point.

In another example, the assembly may include a pin similar to the pin of FIG. 4, with a protrusion added in the pocket surface near the suction side, similar to that described in FIG. 12. In this case, the protrusion may provide additional sealing at the suction side, while the pin, the hole and the airfoil groove still provide at least radial movement of the second end during operation.

In yet another example, the clips of fig. 6 and 7 may be used with the same airfoil second end. In this case, the clip of fig. 7 may be used on one side (e.g., the pressure side) while the spring clip of fig. 6 may be used on the opposite side (e.g., the suction side) to provide controlled movement of the second end during operation.

The above aspects provide various benefits. It will be appreciated that the use of grooves containing support structures or other features may provide a sealing or damping function for the airfoil. The grooves in the inner or outer bands may allow at least one degree of freedom of movement of the airfoil within the grooves, which may reduce operational stresses on the airfoil and extend component life. The various support structures described herein may provide sealing of the grooves, prevent unwanted hot gas ingestion into the grooves, and provide various methods of controlling or damping any excessive movement of the airfoils within the grooves. A resilient or spring-like support structure may be provided for inhibiting the airfoil from moving while still providing freedom of movement. The breakaway support structure may provide an almost complete seal with a smaller gap than current manufacturing methods may produce, while still allowing freedom of movement of the airfoil in the groove. Clips inserted through the holes may utilize known, readily available or mass produced hardware that is easy to install while still providing clearance sealing and controlling the freedom of movement of the airfoil. The use of additive manufactured or printed internal features (e.g., curved dampers) can provide for on-site "assembly" with a monolithic body without the need for additional installation steps to provide a support structure, which can improve process efficiency.

It should be understood that the application of the disclosed design is not limited to turbine engines having fan and booster sections, but is also applicable to turbojet and turboshaft engines.

To the extent not already described, the different features and structures of the various embodiments may be used in combination with each other or instead as desired. One feature is not shown in all embodiments, and is not meant to be so shown, but for brevity of description. Thus, the various features of the different embodiments can be mixed and matched as desired to form new embodiments, whether or not they are explicitly described. All combinations or permutations of features described herein are covered by this disclosure.

This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.

Other aspects of the invention are provided by the subject matter of the following clauses:

1. an assembly for a turbine engine, comprising: a monolithic body comprising: an inner band and an outer band radially spaced from the inner band; a pocket surface defining a groove in one of the inner or outer bands; and an airfoil including an outer wall defining a pressure side and a suction side, a first end extending radially from the other of the inner or outer bands, and a second end extending radially into the groove to define a gap between the airfoil and the one of the inner or outer bands.

2. The assembly of any preceding item, further comprising a support structure at least partially closing the gap and coupled to at least one of the second end or the pocket surface.

3. The assembly of any preceding claim, wherein the support structure comprises: an airfoil recess in the second end; and a pin extending to one of the inner or outer bands and received in the airfoil groove to at least partially close the gap.

4. The assembly of any preceding claim, wherein the support structure comprises a clip spanning the gap and having a first clip end retained by the second end of the airfoil.

5. The assembly of any preceding claim, wherein the clip further comprises a spring clip having a second clip end retained by one of the inner or outer bands.

6. The assembly of any preceding claim, wherein the second end extends completely through one of the inner band or the outer band.

7. The assembly of any preceding claim, wherein the clip further comprises a T-shaped body having a first leg terminating at a first clip end and extending to the second end of the airfoil.

8. The assembly of any preceding claim, wherein the T-shaped body includes a second leg extending perpendicular to the first leg and coupled to one of the inner band or the outer band.

9. The assembly of any preceding claim, further comprising a labyrinth seal at least partially defined by the gap and the first leg of the T-shaped body.

10. The assembly of any preceding item, wherein the support structure includes at least one curvilinear damper connecting the second end of the airfoil and one of the inner band or the outer band.

11. The assembly according to any preceding claim, wherein the support structure extends along the entire axial length of the airfoil.

12. The assembly according to any preceding claim, wherein the support structure is coupled to at least one of the second end or the pocket surface by at least one coupling point.

13. The assembly of any preceding item, wherein the support structure is configured to disengage at the at least one coupling point to define a narrow gap between the second end and the pocket surface that is less than the gap.

14. The assembly of any preceding item, wherein the support structure includes a protrusion defined by the pocket surface and in contact with the second end of the airfoil.

15. An assembly according to any preceding item, wherein the support structure is in contact with the suction side of the airfoil.

16. The assembly of any preceding claim, wherein the support structure completely closes the gap.

17. An assembly according to any preceding claim, wherein the airfoil comprises a static vane in one of a compressor section or a turbine section of the turbine engine.

18. The assembly of any preceding item, wherein the monolithic body further comprises vanes extending completely between the inner band and the outer band.

19. The assembly according to any preceding claim, wherein the monolithic body further includes a pair of airfoils having a first end extending radially from the other of the inner or outer bands and a second end extending radially into the pair of grooves, and a corresponding pair of grooves in one of the inner or outer bands, the vanes being positioned circumferentially between each of the pair of airfoils.

20. A turbine engine, comprising:

an axially arranged compressor, combustor and turbine; and

an assembly in at least one of a compressor or a turbine, the assembly having a monolithic body comprising:

an inner band and an outer band radially spaced from the inner band;

a pocket surface defining a groove in one of the inner or outer bands; and

an airfoil including an outer wall defining a pressure side and a suction side, a first end extending radially from the other of the inner or outer bands, and a second end extending radially into the groove to define a gap between the airfoil and the one of the inner or outer bands.

21. A turbine engine in accordance with any of the preceding claims, further comprising a support structure at least partially closing the gap, wherein the support structure comprises one of a pin, a clip, a curved damper, or a protrusion.

22. An assembly for a turbine engine, the assembly comprising: a monolithic body having an inner band, an outer band radially spaced from the inner band, a groove in one of the inner and outer bands, and a vane having a root at the other of the inner or outer bands and a second tip located within the groove and unaffected by the one of the inner or outer bands.