U.S. patent application number 14/616208 was filed with the patent office on 2016-08-11 for rotor disk sealing and blade attachments system.
This patent application is currently assigned to United Technologies Corporation. The applicant listed for this patent is United Technologies Corporation. Invention is credited to James P. Chrisikos, Brian J. Schwartz.
Application Number | 20160230579 14/616208 |
Document ID | / |
Family ID | 56566642 |
Filed Date | 2016-08-11 |
United States Patent
Application |
20160230579 |
Kind Code |
A1 |
Schwartz; Brian J. ; et
al. |
August 11, 2016 |
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.
Inventors: |
Schwartz; Brian J.; (West
Hartford, CT) ; Chrisikos; James P.; (Vernon,
CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation |
Hartford |
CT |
US |
|
|
Assignee: |
United Technologies
Corporation
Hartford
CT
|
Family ID: |
56566642 |
Appl. No.: |
14/616208 |
Filed: |
February 6, 2015 |
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F01D 5/326 20130101;
F01D 25/04 20130101; F01D 5/3007 20130101; F01D 11/005 20130101;
F01D 5/303 20130101; F01D 11/001 20130101 |
International
Class: |
F01D 11/00 20060101
F01D011/00; F01D 5/30 20060101 F01D005/30 |
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.
Description
FIELD OF INVENTION
[0001] The present disclosure relates to gas turbine engines, and,
more specifically, to a rotor disk with integrated sealing and
blade retention.
BACKGROUND
[0002] 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
[0003] 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.
[0004] 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.
[0005] 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.
[0006] 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.
[0007] 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.
[0008] 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.
[0009] 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
[0010] 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.
[0011] FIG. 1 illustrates an exemplary gas turbine engine, in
accordance with various embodiments;
[0012] FIG. 2 illustrates a sealing system with rotating seal
features formed integral with rotor disks, in accordance with
various embodiments;
[0013] FIG. 3 illustrates a sealing system with a damper and brush
seal between rotor disks, in accordance with various
embodiments;
[0014] 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;
[0015] 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;
[0016] FIG. 4C illustrates a top view of a rotor disk comprising a
retention slot with round corners, in accordance with various
embodiments;
[0017] 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
[0018] 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
[0019] 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.
[0020] 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.
[0021] 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.
[0022] 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.
[0023] 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.
[0024] 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.
[0025] 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.
[0026] 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.
[0027] 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.
[0028] 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.
[0029] 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.
[0030] 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.
[0031] 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.
[0032] 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.
[0033] 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.
[0034] 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.
[0035] 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.
[0036] 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.
[0037] 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.
* * * * *