ROTOR DISK SEALING AND BLADE ATTACHMENTS SYSTEM

Abstract

A rotor disk assembly comprises a circular body configured to rotate about an axis, a contoured slot formed partially through the circular body in an axial direction, and a protrusion extending radially from the circular body adjacent the contoured slot. A turbine or compressor assembly is also provided. The turbine or compressor assembly may include a first disk configured to rotate about an axis, a first contoured slot formed partially through the first disk, a first protrusion adjacent to the first contoured slot and extending radially outward from the first disk, and a first blade disposed in the first contoured slot and configured to engage the first protrusion.

Description

FIELD OF INVENTION

The present disclosure relates to gas turbine engines, and, more specifically, to a rotor disk with integrated sealing and blade retention.

BACKGROUND

Gas turbine engines typically have alternating sets of rotors and stators in the compressor and turbine sections. The rotors may be disks that rotate adjacent to the stators. Sealing between the rotating rotors and the static stators may prevent gas-path air from moving between stages of a compressor or turbine outside of the gas path. A cover plate disposed on the rotating disks may provide sealing. The cover plate may be made separate from the rotor disk and disposed over the rotor disk. The cover plate may also lock a blade into the rotor disk. Adding a cover plate to each rotor in a turbine or compressor may increase the weight and cost of a turbine or compressor section, respectively.

SUMMARY

A rotor disk assembly comprises a circular body configured to rotate about an axis, a contoured slot formed partially through the circular body in an axial direction, and a protrusion extending radially from the circular body adjacent the contoured slot.

In various embodiments, the rotor disk assembly may further comprise a seal disposed on the protrusion. A rotating seal feature may extend from the circular body. The contoured slot may include squared edges. One of the squared edges may be parallel to a radial surface of the protrusion. A blade may be retained in the contoured slot. The blade may engage the protrusion to retain the blade axially within the contoured slot.

A turbine or compressor assembly is also provided. The turbine or compressor assembly may include a first disk configured to rotate about an axis, a first contoured slot formed partially through the first disk, a first protrusion adjacent to the first contoured slot and extending radially outward from the first disk, and a first blade disposed in the first contoured slot and configured to engage the first protrusion.

In various embodiments, the turbine or compressor assembly may further comprise a second disk aft of the first disk, a stator axially between the first disk and the second disk, and a brush seal extending radially inward from the stator. A landing may be coupled between the first disk and the second disk. The brush seal may extend toward the landing. A damper may be coupled between the stator and the brush seal. A second disk may be aft of the first disk, a stator may be axially between the first disk and the second disk, and a first knife seal may extend aft from the first disk towards an interface surface of the stator. A second knife seal may extend forward from the second disk towards the interface surface of the stator. The interface surface of the stator may include a honeycomb configured to deform in response to contact with the first knife seal and/or the second knife seal. The second disk may also include a second contoured slot formed partially through the second disk, a second protrusion adjacent to the first contoured slot and extending radially outward from the second disk, and a second blade disposed in the second contoured slot and configured to engage the second protrusion. The first protrusion may be aft of the first contoured slot and the second protrusion may be aft of the second contoured slot.

A disk sealing system is provided. The disk sealing system comprises a first disk including a first slot and a first protrusion configured to interface with a first blade, and a stator aft of the first disk.

In various embodiments, a first rotating seal feature extends aft from the first disk. The first rotating seal feature may have an annular shape. The stator may further comprise an interface surface and the rotating seal feature may contact the interface surface. The interface surface may comprise a honeycomb. A damper may extend radially inward from the stator and a seal may be disposed at an end of the damper.

The foregoing features and elements may be combined in various combinations without exclusivity, unless expressly indicated otherwise. These features and elements as well as the operation thereof will become more apparent in light of the following description and the accompanying drawings. It should be understood, however, the following description and drawings are intended to be exemplary in nature and non-limiting.

BRIEF DESCRIPTION OF THE DRAWINGS

The subject matter of the present disclosure is particularly pointed out and distinctly claimed in the concluding portion of the specification. A more complete understanding of the present disclosure, however, may best be obtained by referring to the detailed description and claims when considered in connection with the figures, wherein like numerals denote like elements.

FIG. 1 illustrates an exemplary gas turbine engine, in accordance with various embodiments;

FIG. 2 illustrates a sealing system with rotating seal features formed integral with rotor disks, in accordance with various embodiments;

FIG. 3 illustrates a sealing system with a damper and brush seal between rotor disks, in accordance with various embodiments;

FIG. 4A illustrates a partial cross section through a rotor disk having a retention slot to retain a blade on the rotor disk, in accordance with various embodiments;

FIG. 4B illustrates a partial cross section through a rotor disk with a retention slot to retain a blade in an axial direction, in accordance with various embodiments;

FIG. 4C illustrates a top view of a rotor disk comprising a retention slot with round corners, in accordance with various embodiments;

FIG. 5A illustrates a partial cross section of a rotor disk assembly having a blade retained in the rotor disk, in accordance with various embodiments; and

FIG. 5B illustrates a rotor disk assembly from forward looking aft and having a blade retained in the rotor disk, in accordance with various embodiments.

DETAILED DESCRIPTION

The detailed description of exemplary embodiments herein makes reference to the accompanying drawings, which show exemplary embodiments by way of illustration. While these exemplary embodiments are described in sufficient detail to enable those skilled in the art to practice the exemplary embodiments of the disclosure, it should be understood that other embodiments may be realized and that logical changes and adaptations in design and construction may be made in accordance with this disclosure and the teachings herein. Thus, the detailed description herein is presented for purposes of illustration only and not limitation. The scope of the disclosure is defined by the appended claims. For example, the steps recited in any of the method or process descriptions may be executed in any order and are not necessarily limited to the order presented.

Furthermore, any reference to singular includes plural embodiments, and any reference to more than one component or step may include a singular embodiment or step. Also, any reference to attached, fixed, connected or the like may include permanent, removable, temporary, partial, full and/or any other possible attachment option. Additionally, any reference to without contact (or similar phrases) may also include reduced contact or minimal contact. Surface shading lines may be used throughout the figures to denote different parts but not necessarily to denote the same or different materials.

As used herein, “aft” refers to the direction associated with the tail (e.g., the back end) of an aircraft, or generally, to the direction of exhaust of the gas turbine. As used herein, “forward” refers to the direction associated with the nose (e.g., the front end) of an aircraft, or generally, to the direction of flight or motion.

As used herein, “distal” refers to the direction radially outward, or generally, away from the axis of rotation of a turbine engine. As used herein, “proximal” refers to a direction radially inward, or generally, towards the axis of rotation of a turbine engine.

In various embodiments, a seal and disk system with retention structure to retain a blade in a disk as well as sealing structure to seal the gas path may eliminate use of cover plates. Sealing structure formed integral with disks may be cheaper and lighter than cover plates. Similarly, a seal and damper extending from a stator to arms extending from the disk may be less expensive and lighter than cover plates. Additionally, a slot formed partially though the disk and aligned with a protrusion may retain a blade in the disk without a cover plate. Thus, the turbine or compressor section housing a disk as described in the present disclosure may be simplified and made lighter than a disk with a cover plate.

Referring to FIG. 1, a gas turbine engine 100 (such as a turbofan gas turbine engine) is illustrated according to various embodiments. Gas turbine engine 100 is disposed about axial centerline axis 120, which may also be referred to as axis of rotation 120. Gas turbine engine 100 may comprise a fan 140, compressor sections 150 and 160, a combustion section 180, and a turbine section 190. Air compressed in compressor sections 150, 160 may be mixed with fuel and burned in combustion section 180 and expanded across turbine section 190. Turbine section 190 may include high-pressure rotors 192 and low-pressure rotors 194, which rotate in response to the expansion. Turbine section 190 may comprise alternating rows of rotary airfoils or blades 196 and static airfoils or vanes 198. A plurality of bearings 115 may support spools in the gas turbine engine 100. FIG. 1 provides a general understanding of the sections in a gas turbine engine, and is not intended to limit the disclosure. The present disclosure may extend to all types of turbine engines, including turbofan gas turbine engines and turbojet engines, for all types of applications.

The forward-aft positions of gas turbine engine 100 lie along axis of rotation 120. For example, fan 140 may be referred to as forward of turbine section 190 and turbine section 190 may be referred to as aft of fan 140. Typically, during operation of gas turbine engine 100, air flows from forward to aft, for example, from fan 140 to turbine section 190. As air flows from fan 140 to the more aft components of gas turbine engine 100, axis of rotation 120 may also generally define the direction of the air stream flow.

With reference to FIG. 2, sealing system 200 is shown with forward rotor disk 202 and aft rotor disk 204. Forward rotor disk 202 may comprise blade platform 206 to support a blade. Aft rotor disk 204 may comprise a blade platform 208 to retain a blade. Stator 210 includes vane 212 and interface surface 218. Rotating seal feature 216 may extend axially from forward rotor disk 202 towards interface surface 218 of stator 210. In various embodiments, rotating seal feature 216 may be a knife edge seal. Rotating seal feature 216 may make contact with interface surface 218. On a “green” run (i.e., first engine start up), rotating seal feature 216 may contact interface surface 218 as rotating seal feature 216 rotates with forward rotor disk 202. Interface surface 218 may be a honeycomb surface and may deform as rotating seal feature 216 contacts interface surface 218.

In various embodiments, a rotating seal feature 214 may also extend forward from aft rotor disk to interface surface 218 of stator 210. Rotating seal feature 214 may make contact with interface surface 218 and contact interface surface 218 on the green run. Interface surface 218 may deform in response to rotating seal feature 214 contacting interface surface 218.

In various embodiments, rotating seal feature 216 and rotating seal feature 214 may be formed integrally with forward rotor disk 202 and aft rotor disk 204, respectively. Thus, rotating seal feature 216 and rotating seal feature 214 along with forward rotor disk 202 and aft rotor disk 204 may be made from a titanium alloy or a high-performance nickel based alloy (e.g., one of the nickel alloys available under the trade name INCONEL). The contour of rotating seal feature 216 and rotating seal feature 214 may be machined by turning. Rotating seal feature 216 and rotating seal feature 214 may be annular in shape with a portion of the rotating seal feature connecting to forward rotor disk 202 or aft rotor disk 204. Rotating seal feature 216 and rotating seal feature 214 may seal turbine cavities from the gas path.

With reference to FIG. 3, a sealing system 240 comprising a seal 254 is shown, in accordance with various embodiments. Stator 250 may have damper 252 and seal 254 extending into the space between forward rotor disk 242 and aft rotor disk 244 and between forward platform 246 and aft platform 248. Damper 252 may function as a seal having an annular wall and interface with seal 254. Seal 254 may extend to landing 256 built onto arms 258 that attach forward rotor disk 242 to aft rotor disk 244. Seal 254 may seal stages of the turbine or compressor from one another. Damper 252 may dampen vibration modes and provide support for seal 254 at an end of damper 252. In various embodiments, seal 254 may be a brush seal, labyrinth seal, or non-contacting compliant seals. If seal 254 is a brush seal, for example, bristles from the brush seal may extend to and contact landing 256. Seal 254 and damper 252 may form an annular seal structure with one a distal portion of damper 252 anchored to stator 250.

With reference to FIG. 4A, a partial cross section of a rotor disk 280 is shown with protrusion 284 to retain a blade. Rotor disk 280 may be integrated into the sealing systems depicted in FIGS. 2 and 3. Rotor disk 280 has a circular body portion 282 with protrusion 284 at a distal end of circular body portion 282. Protrusion 284 may extend radially outward from rotor disk 280. The distal end of rotor disk 280 has an axial length D1. Protrusion 284 of rotor disk 280 has an axial length D2. The ratio of D1/D2 may be determined by structural requirements of different applications. In various embodiments, the ratio of D1 to D2 may be in the range from two to eight. Circumferential surface 286 of rotor disk 280 may serve as an interface surface for a blade to be attached to a distal end of rotor disk 280. Radial surface 288 defined by a boundary of protrusion 284 may include a seal 290. Seal 290 may be disposed between a later installed blade (i.e., installed on rotor disk 280) and a surface of rotor disk 280 to seal cooling air. The blade may be installed in contoured slot 300, shown by ghosted lines.

With reference to FIG. 4B, rotor disk 280 viewed in the direction from a high pressure side to a low pressure side (forward to aft in a turbine or aft to forward in a compressor) is shown, in accordance with various embodiments. Rotor disk 280 comprises a contoured slot 300 to interface with a turbine blade. Protrusion 284 extends above circumferential surface 286. Seal 290 in radial surface 288 of protrusion 284 is configured to interface with a blade in rotor disk 280.

With reference to FIG. 4C, a top view of contoured slot 300 in rotor disk 280 is shown, in accordance with various embodiments. Contoured slot 300 extends partially through rotor disk 280. Protrusion 284 at a low pressure side of rotor disk 280 may retain a blade in contoured slot 300. Contoured slot 300 may be formed with a contoured disk that leaves rounded edges 310 in contoured slot 300. Contoured slot 300 may be adjacent to protrusion 284 so that a line extending from contoured slot 300 at the aft most point of contoured slot 300 may be coplanar with radial surface 288 of protrusion 284. Rounded edges may be removed or left in place depending on the desired shape of the blade to be retained in contoured slot 300. Upon removing rounded edges, contoured slot 300 may have squared edges 312 and 314 with squared edge 314 parallel to radial surface 288 of protrusion 284.

In various embodiments, contoured slot 300 may be formed using electrochemical machining (ECM), electrical discharge machining (EDM), and/or super abrasive machining (SAM). Contoured slot 300 may also be formed using conventional milling techniques. In various embodiments, SAM is carried out using a grind wheel having a contour similar to the contour of contoured slot 300 (as shown in FIG. 4B). EDM or ECM may be used to remove rounded edges 310 as desired.

In various embodiments, and with reference to FIGS. 5A and 5B, a blade 320 is shown installed in rotor disk 280. Blade 320 may have surface 322 to interface with radial surface 288 and seal 290. Blade platform 324 may extend axially from protrusion 284. Blade 320 may also include surface 326 to rest on and interface with circumferential surface 286 of rotor disk 280. Blade 320 may extend into contoured slot 300 (FIG. 4B) with the surface of blade 320 having a contour matched to contoured slot 300. Contoured slot 300 and protrusion 284 may retain blade 320 axially in rotor disk 280 during use without requiring a contour plate or other extra component to retain blade 320. Protrusion 284 may be on a low pressure side of blade 320 so that the pressure differential between a high pressure side and low pressure side of blade 320 tends to force blade 320 into protrusion 284.

Benefits and other advantages have been described herein with regard to specific embodiments. Furthermore, the connecting lines shown in the various figures contained herein are intended to represent exemplary functional relationships and/or physical couplings between the various elements. It should be noted that many alternative or additional functional relationships or physical connections may be present in a practical system. However, the benefits, advantages, and any elements that may cause any benefit or advantage to occur or become more pronounced are not to be construed as critical, required, or essential features or elements of the disclosure. The scope of the disclosure is accordingly to be limited by nothing other than the appended claims, in which reference to an element in the singular is not intended to mean “one and only one” unless explicitly so stated, but rather “one or more.” Moreover, where a phrase similar to “at least one of A, B, or C” is used in the claims, it is intended that the phrase be interpreted to mean that A alone may be present in an embodiment, B alone may be present in an embodiment, C alone may be present in an embodiment, or that any combination of the elements A, B and C may be present in a single embodiment; for example, A and B, A and C, B and C, or A and B and C.

Systems, methods and apparatus are provided herein. In the detailed description herein, references to “various embodiments”, “one embodiment”, “an embodiment”, “an example embodiment”, etc., indicate that the embodiment described may include a particular feature, structure, or characteristic, but every embodiment may not necessarily include the particular feature, structure, or characteristic. Moreover, such phrases are not necessarily referring to the same embodiment. Further, when a particular feature, structure, or characteristic is described in connection with an embodiment, it is submitted that it is within the knowledge of one skilled in the art to affect such feature, structure, or characteristic in connection with other embodiments whether or not explicitly described. After reading the description, it will be apparent to one skilled in the relevant art(s) how to implement the disclosure in alternative embodiments.

Furthermore, no element, component, or method step in the present disclosure is intended to be dedicated to the public regardless of whether the element, component, or method step is explicitly recited in the claims. No claim element herein is to be construed under the provisions of 35 U.S.C. 112(f), unless the element is expressly recited using the phrase “means for.” As used herein, the terms “comprises”, “comprising”, or any other variation thereof, are intended to cover a non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements does not include only those elements but may include other elements not expressly listed or inherent to such process, method, article, or apparatus.

Claims

1. A rotor disk assembly, comprising:

a circular body configured to rotate about an axis;

a contoured slot formed partially through the circular body in an axial direction; and

a protrusion extending radially from the axis of the circular body adjacent the contoured slot.

2. The rotor disk assembly of claim 1, further comprising a seal disposed on the protrusion.

3. The rotor disk assembly of claim 1, further comprising a rotating seal feature extending from the circular body.

4. The rotor disk assembly of claim 1, wherein the contoured slot comprises squared edges.

5. The rotor disk assembly of claim 4, wherein at least one of the squared edges is parallel to a radial surface of the protrusion.

6. The rotor disk assembly of claim 1, further comprising a blade retained in the contoured slot.

7. The rotor disk assembly of claim 6, wherein the blade engages the protrusion to retain the blade axially within the contoured slot.

8. A turbine assembly, comprising:

a first disk configured to rotate about an axis;

a first contoured slot formed partially through the first disk;

a first protrusion adjacent to the first contoured slot and extending radially outward from the first disk; and

a first blade disposed in the first contoured slot and configured to engage the first protrusion.

9. The turbine assembly of claim 8, further comprising:

a second disk aft of the first disk;

a stator axially between the first disk and the second disk; and

a brush seal extending radially inward from the stator.

10. The turbine assembly of claim 9, further comprising a landing coupled between the first disk and the second disk, wherein the brush seal extends toward the landing.

11. The turbine assembly of claim 10, further comprising a damper coupled between the stator and the brush seal.

12. The turbine assembly of claim 8, further comprising:

a second disk aft of the first disk;

a stator axially between the first disk and the second disk;

a first knife seal extending aft from the first disk towards an interface surface of the stator; and

a second knife seal extending forward from the second disk towards the interface surface of the stator.

13. The turbine assembly of claim 12, wherein the interface surface of the stator comprises a honeycomb structure configured to deform in response to contact with at least one of the first knife seal and the second knife seal.

14. The turbine assembly of claim 12, wherein the second disk comprises:

a second contoured slot formed partially through the second disk;

a second protrusion adjacent to the first contoured slot and extending radially outward from the second disk; and

a second blade disposed in the second contoured slot and configured to engage the second protrusion.

15. The turbine assembly of claim 14, wherein the first protrusion is aft of the first contoured slot and the second protrusion is aft of the second contoured slot.

16. A disk sealing system, comprising:

a first disk including a first slot and a first protrusion configured to interface with a first blade;

a stator aft of the first disk.

17. The disk sealing system of claim 16, further comprising a first rotating seal feature extending aft from the first disk, the first rotating seal feature having an annular shape.

18. The disk sealing system of claim 16, wherein the stator further comprises an interface surface, wherein the rotating seal feature contacts the interface surface.

19. The disk sealing system of claim 18, wherein the interface surface comprises a honeycomb structure.

20. The disk sealing system of claim 16, further comprising:

a damper extending radially inward from the stator; and

a seal at an end of the damper.