U.S. patent application number 15/063089 was filed with the patent office on 2017-09-07 for airfoil tip geometry to reduce blade wear in gas turbine engines.
The applicant listed for this patent is General Electric Company. Invention is credited to Ananda Barua, Kenneth Martin Lewis, Sathyanarayanan Raghavan, Neelesh Nandkumar Sarawate, Changjie Sun.
Application Number | 20170254340 15/063089 |
Document ID | / |
Family ID | 59723483 |
Filed Date | 2017-09-07 |
United States Patent
Application |
20170254340 |
Kind Code |
A1 |
Raghavan; Sathyanarayanan ;
et al. |
September 7, 2017 |
AIRFOIL TIP GEOMETRY TO REDUCE BLADE WEAR IN GAS TURBINE
ENGINES
Abstract
An airfoil for use in a turbomachine includes a pressure
sidewall and a suction sidewall coupled to the pressure sidewall.
The suction sidewall and the pressure sidewall define a leading
edge and an opposite trailing edge. The leading edge and the
trailing edge define a chord distance. The airfoil further includes
a root portion, and a tip portion. The tip portion extends between
the pressure sidewall and the suction sidewall such that the tip
portion is substantially perpendicular to each sidewall. The tip
portion includes at least one planar section and at least one
oblique section that forms a recess within the tip portion. The at
least one oblique section extends from the at least one planar
section towards the root portion along the chord distance. The tip
portion is configured to reduce airfoil wear during contact with a
surrounding casing.
Inventors: |
Raghavan; Sathyanarayanan;
(Ballston Lake, NY) ; Sarawate; Neelesh Nandkumar;
(Niskayuna, NY) ; Lewis; Kenneth Martin; (Liberty
Township, OH) ; Sun; Changjie; (Clifton Park, NY)
; Barua; Ananda; (Schenectady, NY) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
General Electric Company |
Schenectady |
NY |
US |
|
|
Family ID: |
59723483 |
Appl. No.: |
15/063089 |
Filed: |
March 7, 2016 |
Current U.S.
Class: |
1/1 |
Current CPC
Class: |
F04D 29/164 20130101;
F04D 29/324 20130101 |
International
Class: |
F04D 29/32 20060101
F04D029/32 |
Claims
1. An airfoil for use in a turbomachine, said airfoil comprising: a
pressure sidewall; a suction sidewall coupled to said pressure
sidewall, wherein said suction sidewall and said pressure sidewall
define a leading edge and an opposite trailing edge, wherein said
leading edge and said trailing edge define a chord distance; a root
portion; and a tip portion extending between said pressure sidewall
and said suction sidewall such that said tip portion is
substantially perpendicular to each sidewall, said tip portion
comprises at least one planar section and at least one oblique
section that forms a recess within said tip portion, said at least
one oblique section extends from said at least one planar section
towards said root portion along said chord distance, said tip
portion is configured to reduce airfoil wear during contact with a
surrounding casing.
2. The airfoil in accordance with claim 1, wherein said at least
one oblique section extends from said leading edge within a range
from approximately 5% to approximately 15% of said chord distance
to said at least one planar section, wherein said leading edge has
a first length that extends between said root portion and said at
least one oblique section and said trailing edge has a second
length that extends between said root portion and said at least one
planar section such that said first length is less than said second
length.
3. The airfoil in accordance with claim 1, wherein said at least
one oblique section extends from said leading edge within a range
from approximately 15% to approximately 30% of said chord distance
to said at least one planar section, wherein said leading edge has
a first length that extends between said root portion and said at
least one oblique section and said trailing edge has a second
length that extends between said root portion and said at least one
planar section such that said first length is less than said second
length.
4. The airfoil in accordance with claim 1, wherein said at least
one oblique section extends from said leading edge within a range
from approximately 30% to approximately 50% of said chord distance
to said at least one planar section, wherein said leading edge has
a first length that extends between said root portion and said at
least one oblique section and said trailing edge has a second
length that extends between said root portion and said at least one
planar section such that said first length is less than said second
length.
5. The airfoil in accordance with claim 1, wherein said at least
one oblique section extends from said trailing edge within a range
from approximately 5% to approximately 15% of said chord distance
to said at least one planar section, wherein said leading edge has
a first length that extends between said root portion and said at
least one planar section and said trailing edge has a second length
that extends between said root portion and said at least one
oblique section such that said first length is greater than said
second length.
6. The airfoil in accordance with claim 1, wherein said at least
one oblique section extends from said trailing edge within a range
from approximately 15% to approximately 30% of said chord distance
to said at least one planar section, wherein said leading edge has
a first length that extends between said root portion and said at
least one planar section and said trailing edge has a second length
that extends between said root portion and said at least one
oblique section such that said first length is greater than said
second length.
7. The airfoil in accordance with claim 1, wherein said at least
one oblique section extends from said trailing edge within a range
from approximately 30% to approximately 50% of said chord distance
to said at least one planar section, wherein said leading edge has
a first length that extends between said root portion and said at
least one planar section and said trailing edge has a second length
that extends between said root portion and said at least one
oblique section such that said first length is greater than said
second length.
8. The airfoil in accordance with claim 1, wherein said at least
one oblique section comprises a first oblique section and a second
oblique section, said first oblique section extends from said
leading edge approximately 15% of said chord distance to said at
least one planar section and said second oblique section extends
from said trailing edge approximately 15% of said chord distance to
said at least one planar section.
9. The airfoil in accordance with claim 8, wherein said leading
edge has a first length that extends between said root portion and
said first oblique section and said trailing edge has a second
length that extends between said root portion and said second
oblique section such that said first length is substantially equal
to said second length.
10. The airfoil in accordance with claim 1, wherein said at least
one oblique section extends from said at least one planer section
towards said root portion within a range including approximately 2
mils to approximately 5 mils.
11. The airfoil in accordance with claim 1, wherein said at least
one planar section comprises a first planar section adjacent said
leading edge and a second planar section adjacent trailing edge,
said at least one oblique section extends between said first planar
section and said second planar section.
12. The airfoil in accordance with claim 1, wherein said at least
one oblique section is defined with a curve.
13. A turbomachine comprising: a casing; a rotor assembly, said
casing at least partially extending about said rotor assembly, said
rotor assembly comprising: a rotor shaft; a plurality of rotor
blades coupled to said rotor shaft, each rotor blade of said
plurality of rotor blades comprises an airfoil comprising a
pressure sidewall and a suction sidewall coupled to said pressure
sidewall, wherein said suction sidewall and said pressure sidewall
define a leading edge and an opposite trailing edge, wherein said
leading edge and said trailing edge define a chord distance, said
airfoil further comprising a root portion and a tip portion
extending between said pressure sidewall and said suction sidewall
such that said tip portion is substantially perpendicular to each
sidewall, said tip portion comprising at least one planar section
and at least one oblique section that forms a recess within said
tip portion, said at least one oblique section extends from said at
least one planar section towards said root portion along said chord
distance, said tip portion is configured to reduce rotor blade wear
during contact with said casing.
14. The turbomachine in accordance with claim 13, wherein said at
least one oblique section extends from said leading edge to said at
least one planar section, wherein a distance measured between said
casing and said leading edge is greater than a distance measured
between said casing and said trailing edge.
15. The turbomachine in accordance with claim 13, wherein said at
least one oblique section extends from said trailing edge to said
at least one planar section, wherein a distance measured between
said casing and said trailing edge is greater than a distance
measured between said casing and said leading edge.
16. The turbomachine in accordance with claim 13, wherein said at
least one oblique section comprises a first oblique section and a
second oblique section, said first oblique section extends from
said leading edge approximately 15% of said chord distance to said
at least one planar section and said second oblique section extends
from said trailing edge approximately 15% of said chord distance to
said at least one planar section.
17. A method for reducing blade wear during turbomachine operation,
the turbomachine including a casing, a rotor shaft, and a plurality
of rotor blades, each rotor blade of the plurality of rotor blades
including an airfoil including a pressure sidewall and a suction
sidewall coupled to the pressure sidewall, wherein the suction
sidewall and the pressure sidewall define a leading edge and an
opposite trailing edge, wherein the leading edge and the trailing
edge define a chord distance, the airfoil further includes a root
portion and a tip portion extending between the pressure sidewall
and the suction sidewall such that the tip portion is substantially
perpendicular to each sidewall, said method comprising: removing
blade material from the tip portion comprising forming a recess
from at least one oblique section adjacent at least one planar
section on the tip portion, the at least one oblique section
extends from the at least one planar section towards the root
portion along the chord distance; and coupling the rotor blade to
the rotor shaft such that during turbomachine operation when the
tip portion contacts the casing, wear of the rotor blade is
reduced.
18. The method in accordance with claim 17, wherein removing blade
material from the tip portion further comprises removing blade
material from the tip portion at the leading edge such that the
leading edge has a first length that extends between the root
portion and the at least one oblique section and the trailing edge
has a second length that extends between the root portion and the
at least one planar section such that the first length is less than
the second length.
19. The method in accordance with claim 17, wherein removing blade
material from the tip portion further comprises removing blade
material from the tip portion at the trailing edge such that the
leading edge has a first length that extends between the root
portion and the at least one planar section and the trailing edge
has a second length that extends between the root portion and the
at least one oblique section such that the first length is greater
than the second length.
20. The method in accordance with claim 17, wherein removing blade
material from the tip portion further comprises: removing the
leading edge tip portion to form a first oblique section; and
removing the trailing edge tip portion to form a second oblique
section.
Description
BACKGROUND
[0001] The field of the disclosure relates generally to gas turbine
engines and, more particularly, to airfoil tip geometry to reduce
blade wear in gas turbine engines.
[0002] At least some known turbomachines, i.e., gas turbine
engines, include a compressor that compresses air through a
plurality of rotatable compressor blades enclosed within a
compressor casing, and a combustor that ignites a fuel-air mixture
to generate combustion gases. The combustion gases are channeled
through rotatable turbine blades in a turbine through a hot gas
path. Such known turbomachines convert thermal energy of the
combustion gas stream to mechanical energy used to generate thrust
and/or rotate a turbine shaft to power an aircraft. Output from the
turbomachine may also be used to power a machine, such as, an
electric generator, a compressor, or a pump.
[0003] Under some known operating conditions, rub events occur
within the turbomachine, wherein a rotor blade tip contacts or rubs
against the surrounding stationary casing inducing radial and
tangential loads into a rotor blade airfoil. Generally during rub
events, these loads cause the rotor blade to vibrate and deflect
causing wear thereto. Excessive tip rub events cause wear to the
rotor blade including, but not limited to, loss of blade material,
which decreases turbomachine performance.
[0004] During tip rub events, the rotor blade is known to lose more
material from the tip than the penetration distance into the
casing. For example, if the blade tip penetrates the casing 1 mil
(25.4 micrometers (.mu.m)) then the blade tip is known to lose as
much as 10 mils (254 .mu.m) of material. The thickness of material
lost in the blade tip divided by the penetration distance into the
casing is known as a rub ratio. In the above example, the rub ratio
would be 10:1, or known to have a rub ratio value of 10.
Turbomachines with a high rub ratio are known to have decreased
performance and decreased service life resulting in higher
maintenance costs.
BRIEF DESCRIPTION
[0005] In one aspect, an airfoil for use in a turbomachine is
provided. The airfoil includes a pressure sidewall and a suction
sidewall coupled to the pressure sidewall, the suction sidewall and
the pressure sidewall define a leading edge and an opposite
trailing edge. The leading edge and the trailing edge define a
chord distance. The airfoil further includes a root portion, and a
tip portion. The tip portion extends between the pressure sidewall
and the suction sidewall such that the tip portion is substantially
perpendicular to each sidewall. The tip portion includes at least
one planar section and at least one oblique section that forms a
recess within the tip portion. The at least one oblique section
extends from the at least one planar section towards the root
portion to along the chord distance. The tip portion is configured
to reduce airfoil wear during contact with a surrounding
casing.
[0006] In a further aspect, a turbomachine is provided. The
turbomachine includes a casing, and a rotor assembly, the casing at
least partially extending about the rotor assembly. The rotor
assembly includes a rotor shaft, and a plurality of rotor blades
coupled to the rotor shaft. Each rotor blade of the plurality of
rotor blades includes an airfoil including a pressure sidewall and
a suction sidewall coupled to the pressure sidewall. The suction
sidewall and the pressure sidewall define a leading edge and an
opposite trailing edge. The leading edge and the trailing edge
define a chord distance. The airfoil further includes a root
portion, and a tip portion. The tip portion extends between the
pressure sidewall and the suction sidewall such that the tip
portion is substantially perpendicular to each sidewall. The tip
portion includes at least one planar section and at least one
oblique section that forms a recess within said tip portion. The at
least one oblique section slopes from the at least one planar
section towards the root portion along the chord distance. The tip
portion is configured to reduce rotor blade wear during contact
with the casing.
[0007] In another aspect, a method for reducing blade wear during
turbomachine operation is provided. The turbomachine includes a
casing, a rotor shaft, and a plurality of rotor blades. Each rotor
blade of the plurality of rotor blades includes an airfoil
including a pressure sidewall and a suction sidewall coupled to the
pressure sidewall. The suction sidewall and the pressure sidewall
define a leading edge and an opposite trailing edge. The leading
edge and the trailing edge define a chord distance. The airfoil
further includes a root portion, and a tip portion. The tip portion
extends between the pressure sidewall and the suction sidewall such
that the tip portion is substantially perpendicular to each
sidewall. The method includes removing blade material from the tip
portion including forming a recess from at least one oblique
section adjacent to at least one planar section on the tip portion.
The at least one oblique section extends from the at least one
planar section towards the root portion along the chord distance.
The method further includes coupling the rotor blade to the rotor
shaft such that during turbomachine operation, when the tip portion
contacts the casing, wear of the rotor blade is reduced.
DRAWINGS
[0008] These and other features, aspects, and advantages of the
present disclosure will become better understood when the following
detailed description is read with reference to the accompanying
drawings in which like characters represent like parts throughout
the drawings, wherein:
[0009] FIG. 1 is a schematic diagram of an exemplary turbomachine,
i.e., a turbofan;
[0010] FIG. 2 is a perspective view of an exemplary rotor blade
that may be used within the turbomachine shown in FIG. 1;
[0011] FIG. 3 is a schematic view of an exemplary tip portion of
the rotor blade shown in FIG. 2;
[0012] FIG. 4 is a graphical view of operational features of the
tip portion shown in FIG. 3;
[0013] FIG. 5 is a schematic view of an alternative tip portion
that may be used with the rotor blade shown in FIG. 2;
[0014] FIG. 6 is a schematic view of another alternative tip
portion that may be used with the rotor blade shown in FIG. 2;
and
[0015] FIG. 7 is a schematic view of a further alternative tip
portion that may be used with the rotor blade shown in FIG. 2.
[0016] Unless otherwise indicated, the drawings provided herein are
meant to illustrate features of embodiments of this disclosure.
These features are believed to be applicable in a wide variety of
systems comprising one or more embodiments of this disclosure. As
such, the drawings are not meant to include all conventional
features known by those of ordinary skill in the art to be required
for the practice of the embodiments disclosed herein.
DETAILED DESCRIPTION
[0017] In the following specification and claims, reference will be
made to a number of terms, which shall be defined to have the
following meanings.
[0018] The singular forms "a", "an", and "the" include plural
references unless the context clearly dictates otherwise.
[0019] "Optional" or "optionally" means that the subsequently
described event or circumstance may or may not occur, and that the
description includes instances where the event occurs and instances
where it does not.
[0020] Approximating language, as used herein throughout the
specification and claims, may be applied to modify any quantitative
representation that could permissibly vary without resulting in a
change in the basic function to which it is related. Accordingly, a
value modified by a term or terms, such as "about",
"approximately", and "substantially", are not to be limited to the
precise value specified. In at least some instances, the
approximating language may correspond to the precision of an
instrument for measuring the value. Here and throughout the
specification and claims, range limitations may be combined and/or
interchanged, such ranges are identified and include all the
sub-ranges contained therein unless context or language indicates
otherwise.
[0021] Rotor blade tip geometries as described herein provide a
method for reducing blade wear in a turbomachine. Specifically, a
rotor blade includes an airfoil having a suction sidewall coupled
to a pressure sidewall at a leading edge and a trailing edge. A tip
portion extends between the suction sidewall and the pressure
sidewall and includes a planar section and an oblique section. In
some embodiments, the tip portion includes a first oblique section
and a second oblique section. Modifying the rotor blade tip
geometry by grinding the tip portion and forming the oblique
section reduces the rub ratio of the rotor blade, and thereby, the
wear of the rotor blade. Specifically, the oblique section is sized
such that a contact area between the rotor blade and a surrounding
casing is reduced, thereby decreasing the radial and tangential
loads induced into the rotor blade during a rub event. Reducing the
loads resulting from a rub event decreases vibration and deflection
of the rotor blade and reduces material loss at the tip portion.
Furthermore, modifying the rotor blade tip geometry changes the
vibratory modes of the rotor blade such that radial elongation is
decreased further reducing material loss at the tip portion.
Additionally, a reduction in radial deflection allows the rotor
blade to be positioned closer to the surrounding casing.
Accordingly, decreasing the rub ratio of the rotor blade decreases
wear and material loss during a rub event, increases turbomachine
performance, and reduces maintenance costs.
[0022] As used herein, the terms "axial", and "axially", refer to
directions and orientations which extend substantially parallel to
a centerline 138, as shown in FIG. 1, of a turbine engine.
Moreover, the terms "radial", and "radially", refer to directions
and orientations which extend substantially perpendicular to
centerline 138 of the turbine engine. In addition, as used herein,
the terms "circumferential", and "circumferentially", refer to
directions and orientations which extend arcuately about centerline
138 of the turbine engine. The term "fluid", as used herein,
includes any medium or material that flows, including, but not
limited to, air.
[0023] FIG. 1 is a schematic view of a turbomachine 100, i.e., a
gas turbine engine, and more specifically, an aircraft engine or
turbofan. In the exemplary embodiment, turbomachine 100 includes an
air intake section 102, and a compressor section 104 that is
coupled downstream from, and in flow communication with, intake
section 102. Compressor section 104 is enclosed within a compressor
casing 106. A combustor section 108 is coupled downstream from, and
in flow communication with, compressor section 104, and a turbine
section 110 is coupled downstream from, and in flow communication
with, combustor section 108. Turbine section 110 is enclosed within
a turbine casing 112 and includes an exhaust section 114 that is
downstream from turbine section 110. A combustor housing 116
extends about combustor section 108 and is coupled to compressor
casing 106 and turbine casing 112. Moreover, in the exemplary
embodiment, turbine section 110 is coupled to compressor section
104 through a rotor assembly 118 that includes, without limitation,
a compressor rotor, or drive shaft 120 and a turbine rotor, or
drive shaft 122.
[0024] In the exemplary embodiment, combustor section 108 includes
a plurality of combustor assemblies, i.e., combustors 124 that are
each coupled in flow communication with compressor section 104.
Combustor section 108 also includes at least one fuel nozzle
assembly 126. Each combustor 108 is in flow communication with at
least one fuel nozzle assembly 126. Moreover, in the exemplary
embodiment, turbine section 110 and compressor section 104 are
rotatably coupled to a fan assembly 128 through drive shaft 120.
Alternatively, turbomachine 100 may be a gas turbine engine and for
example, and without limitation, be rotatably coupled to an
electrical generator and/or a mechanical drive application, e.g., a
pump. In the exemplary embodiment, compressor section 104 includes
at least one compressor stage that includes a compressor blade
assembly 130 and an adjacent stationary stator vane assembly 132.
Each compressor blade assembly 130 includes a plurality of
circumferentially spaced blades (not shown) and is coupled to rotor
assembly 118, or, more specifically, compressor drive shaft 120.
Each stator vane assembly 132 includes a plurality of
circumferentially spaced stator vanes (not shown) and is coupled to
compressor casing 106. Also, in the exemplary embodiment, turbine
section 110 includes at least one turbine blade assembly 134 and at
least one adjacent stationary nozzle assembly 136. Each turbine
blade assembly 134 is coupled to rotor assembly 118, or, more
specifically, turbine drive shaft 122 along a centerline 138.
[0025] In operation, air intake section 102 channels air 140
towards compressor section 104. Compressor section 104 compresses
air 140 to higher pressures and temperatures prior to discharging
compressed air 142 towards combustor section 108. Compressed air
142 is channeled to fuel nozzle assembly 126, mixed with fuel (not
shown), and burned within each combustor 124 to generate combustion
gases 144 that are channeled downstream towards turbine section
110. After impinging turbine blade assembly 134, thermal energy is
converted to mechanical rotational energy that is used to drive
rotor assembly 118. Turbine section 110 drives compressor section
104 and/or fan assembly 128 through drive shafts 120 and 122, and
exhaust gases 146 are discharged through exhaust section 114 to the
ambient atmosphere.
[0026] FIG. 2 is a perspective view of an exemplary rotor blade
200, and more specifically, a compressor blade, that may be found
within turbomachine 100 (shown in FIG. 1). In the exemplary
embodiment, rotor blade 200 includes an airfoil 202, a platform
204, and a dovetail 206 that is used for mounting rotor blade 200
to compressor drive shaft 120 (shown in FIG. 1). Airfoil 202
includes a root portion 208, adjacent platform 204, and an opposite
tip portion 210. Further, airfoil 202 includes a pressure sidewall
212 and an opposite suction sidewall 214. In the exemplary
embodiment, pressure sidewall 212 is substantially concave and
suction sidewall 214 is substantially convex. Pressure sidewall 212
is coupled to suction sidewall 214 at a leading edge 216 and at an
axially spaced trailing edge 218. Trailing edge 218 is spaced
chord-wise and downstream from leading edge 216. Pressure sidewall
212 and suction sidewall 214 each extend longitudinally or radially
outward in a length 220 from root portion 208 to blade tip portion
210. Along a chord of blade 200, a mid-chord line 217 is defined at
the mid-point of the chord. Tip portion 210 is defined between
sidewalls 212 and 214 and includes a planar section 222 that is
defined as the radially outer surface of blade 200 and
substantially perpendicular to each sidewall 212 and 214. Tip
portion 210 also includes an oblique section 300 adjacent to planar
section 222 and described further below in reference to FIG. 3. In
an alternative embodiment, rotor blade 200 may have any other
configuration that enables turbomachine to function as described
herein.
[0027] In the exemplary embodiment, compressor casing 106
circumferentially extends around rotor blade 200, and tip portion
210. Specifically, tip portion 210 at leading edge 216 and oblique
section 300 has a gap distance 224 that is substantially not equal
to a gap distance 226 of tip portion 210 at trailing edge 218 and
planar section 222. Furthermore, a flow path 228 for compressed air
142 (shown in FIG. 1) is defined between compressor casing 106 and
shaft 120.
[0028] During operation, rotor blade 200 rotates within casing 106
about centerline 138 (shown in FIG. 1). In some operating
conditions, such as an imbalanced load, rotor blade 200,
specifically tip portion 210, contacts or rubs against casing 106,
which is also known as a rub event. Specifically, tip portion 210
is jammed into casing 106, such that radial and tangential loads
are induced into rotor blade 200. Generally during rub events,
these loads cause rotor blade 200 to vibrate and deflect causing
wear thereto. The deflection of rotor blade 200, at least in part,
depends on the vibratory modes of the blade that are excited during
the rub event. Some vibratory modes are known to increase radial
elongation of rotor blade 200 resulting in an increased amount of
wear to tip portion 210.
[0029] At least some of the wear rotor blade 200 incurs during the
rub event includes material loss from tip portion 210.
Specifically, when tip portion 210 contacts casing 106, rotor blade
200 loses material at tip portion 210 such that overall length 220
is reduced. A rub ratio is a value that may be used to quantify the
amount of wear rotor blade 200 experiences during the rub event. A
rub ratio is defined as a thickness of material lost from tip
portion 210 during a rub event divided by an amount of penetration
by tip portion 210 into casing 106. For example, if tip portion 210
penetrates into the casing 1 mil (25 .mu.m) and 10 mils (101 .mu.m)
of blade material is lost from tip portion 210, the rub ratio is
10.
[0030] FIG. 3 is a schematic view of an exemplary tip portion 210
for use with rotor blade 200. In the exemplary embodiment, tip
portion 210 includes planar section 222 that extends from pressure
sidewall 212 to suction sidewall 214 and substantially
perpendicular thereto. Additionally, tip portion 210 includes an
oblique section 300 that slopes from planar section 222 inwards
towards root portion 208 to leading edge 216 forming a recess 301.
Oblique section 300 also extends from pressure sidewall 212 to
suction sidewall 214 and is substantially perpendicular thereto. In
the exemplary embodiment, oblique section 300 extends a distance
302 along tip portion 210. Specifically, oblique section 300
extends along tip portion 210 from leading edge 216 within a range
from approximately 5% to approximately 50% of a chord distance 304
of airfoil 202. For example, oblique section 300 extends along tip
portion 210 from leading edge 216 within a range from approximately
5% to approximately 15% of a chord distance 304 of airfoil 202.
More specifically, in the illustrated embodiment, oblique section
300 extends along tip portion 210 from leading edge 216
approximately 15% of chord distance 304. Oblique section 300 also
has a depth 306 from planar section 222 such that a length 308 of
leading edge 216 that extends from tip portion 210 to root portion
208 is shorter than a length 310 of trailing edge 218 from tip
portion 210 to root portion 208. Said another way, distance 224
(shown in FIG. 2) between casing 106 (shown in FIG. 2) and leading
edge 216 is greater than distance 226 (shown in FIG. 2) between
casing 106 and trailing edge 218. In the exemplary embodiment,
depth 306 is within a range including approximately 2 mils (51
.mu.m) to approximately 5 mils (127 .mu.m). In alternative
embodiments, depth 306 may have any other distance that enables tip
portion 210 to function as described herein.
[0031] In some embodiments, for example, oblique section 300 is
formed at line 312 that extends a distance 314 along tip portion
210 from leading edge 216 within a range from approximately 15% to
approximately 30% of chord distance 304 forming recess 301.
Specifically, in the illustrated embodiment, oblique section line
312 extends approximately 30% of chord distance 304 from leading
edge 216. Extending recess 301 further from leading edge 216, such
as with oblique section line 312, reduces the area of planar
section 222 that contacts with casing 106 during a rub event
thereby lowering the contact force between rotor blade 200 and
casing 106. In other embodiments, for example, oblique section 300
is formed at line 316 that extends a distance 318 along tip portion
210 from leading edge 216 within a range from approximately 30% to
approximately 50% of chord distance 304 forming recess 301.
Specifically, in the illustrated embodiment, oblique section line
316 extends approximately 50% of chord distance 304 from leading
edge 216. Extending recess 301 further from leading edge 216, such
as with oblique section line 316, further reduces the area of
planar section 222 that contacts with casing 106 during a rub event
thereby lowering the contact force between rotor blade 200 and
casing 106. In further embodiments, oblique section 300 may extend
any other distance along tip portion 210 from leading edge 216 that
enables tip portion 210 to function as described herein.
[0032] Additionally, in some embodiments, an oblique section 320 is
defined from trailing edge 218 such that a length of trailing edge
218 from tip portion 210 to root portion 208 is shorter than a
length of leading edge from tip portion 210 to root portion 208.
Said another way, distance 226 between casing 106 and trailing edge
218 is greater than distance 224 between casing 106 and leading
edge 216. In the exemplary embodiment, oblique section 320 extends
along tip portion 210 from trailing edge 218 within a range from
approximately 5% to approximately 50% of chord distance 304 of
airfoil 202. For example, oblique section 320 extends a distance
322 from trailing edge 218 within a range from approximately 5% to
approximately 15% of chord distance 304 forming a recess 321.
Specifically, in the illustrated embodiment, oblique section 320
extends approximately 15% of chord distance 304 from trailing edge
218. In other embodiments, for example, oblique section 320 is
formed at line 324 that extends a distance 326 from trailing edge
218 within a range from approximately 15% to approximately 30% of
chord distance 304 forming recess 321. Specifically, in the
illustrated embodiment, oblique section line 324 extends
approximately 30% of chord distance 304 from trailing edge 218.
Extending recess 321 further from trailing edge 216, such as with
oblique section line 324, reduces the area of planar section 222
that contacts with casing 106 during a rub event thereby lowering
the contact force between rotor blade 200 and casing 106. In yet
other embodiments, for example, oblique section 320 is formed at
line 328 that extends a distance 330 from trailing edge 218 within
a range from approximately 30% to approximately 50% of chord
distance 304 forming recess 321. Specifically, in the illustrated
embodiment, oblique section line 328 extends approximately 50% of
chord distance 304 from trailing edge 218. Extending recess 321
further from trailing edge 218, such as with oblique section line
328, reduces the area of planar section 222 that contacts with
casing 106 during a rub event thereby lowering the contact force
between rotor blade 200 and casing 106. In alternative embodiments,
oblique section 320 extends any other distance along tip portion
210 from trailing edge 218 that enables tip portion 210 to function
as described herein.
[0033] Furthermore, in some embodiments, tip portion 210 includes
oblique sections on both leading edge 216 and trailing edge 218.
For example, tip portion 210 includes oblique section 300 and
oblique section 320 such that a length of leading edge 216 from tip
portion 210 to root portion 208 is substantially equal to a length
of trailing edge 218 from tip portion 210 to root portion 208. Said
another way, distance 224 between casing 106 and leading edge 216
is substantially equal to distance 226 between casing 106 and
trailing edge 218.
[0034] In the exemplary embodiment, oblique section 300 is formed
by grinding tip portion 210 and removing rotor blade 200 material
in a machine shop using known machining techniques. Alternatively,
oblique section 300 can be formed by any other method that enables
rotor blade 200 to function as described herein.
[0035] FIG. 4 is a graphical view, i.e., chart 400, of operational
features of tip portion 210 shown in FIGS. 2-3. Specifically, chart
400 illustrates a rub ratio value for four different tip geometries
of tip portion 210 (shown in FIG. 3). The rub ratio is defined as a
thickness of material lost from tip portion 210 during a rub event
divided by an amount of penetration by tip portion 210 into casing
106 as described in reference to FIG. 2. Chart 400 includes a
y-axis 402 defining the rub ratio value on a linear scale. Along
the x-axis, four different tip geometries are shown: a baseline
geometry 404, which includes planar section 222 (shown in FIG. 3)
that extends the full length of tip portion 210 from leading edge
216 (shown in FIG. 3) to trailing edge 218 (shown in FIG. 3); a
first geometry 406, which includes oblique section 300 (shown in
FIG. 3) adjacent to leading edge 216; a second geometry 408, which
includes oblique section 320 (shown in FIG. 3) adjacent to trailing
edge 218; and a third geometry 410, which includes both oblique
sections 300 and 320.
[0036] In the exemplary chart 400, each tip geometry 404, 406, 408,
and 410 is subjected to a rub event with casing 106 (shown in FIG.
1) and a thickness of material loss at each of leading edge 216,
mid-chord line 217 (shown in FIG. 3), and trailing edge 218 are
recorded. Then the rub ratio at each leading edge 216, mid-chord
line 217, and trailing edge 218 are determined. Chart 400 includes
a first group of bars 412 that represents the rub ratio for tip
portion 210 with baseline geometry 404. A leftmost bar 414
represents the rub ratio at leading edge 216 of baseline geometry
404, a middle bar 416 represents the rub ratio at mid-chord line
217, and a rightmost bar 418 represents the rub ratio at trailing
edge 218.
[0037] Further, in the exemplary chart 400, a second group of bars
420 represents the rub ratio for tip portion 210 with first tip
geometry 406. A leftmost bar 422 represents the rub ratio at
leading edge 216 which is less than the rub ratio of baseline
geometry 404 thereby reducing wear to tip portion 210 during a rub
event. A middle bar 424 represents the rub ratio at mid-chord line
217, and a rightmost bar 426 represents the rub ratio at trailing
edge 218.
[0038] A third group of bars 428 represents the rub ratio for tip
portion 210 with second tip geometry 408. A leftmost bar 430
represents the rub ratio at leading edge 216, a middle bar 432
represents the rub ratio at mid-chord line 217, and a rightmost bar
434 represents the rub ratio at trailing edge 218. At each
location, leading edge 216, mid-chord line 217, and trailing edge
218, the rub ratio is less than baseline geometry 404 thereby
reducing wear of tip portion 210 during a rub event.
[0039] A fourth group of bars 436 represents the rub ratio for tip
portion 210 with third tip geometry 410. A leftmost bar 438
represents the rub ratio at leading edge 216, a middle bar 440
represents the rub ratio at mid-chord line 217, and a rightmost bar
442 represents the rub ratio at trailing edge 218. At each
location, leading edge 216, mid-chord line 217, and trailing edge
218, the rub ratio is lower than baseline geometry 404 thereby
reducing wear of tip portion 210 during a rub event.
[0040] As shown in chart 400, modifying the geometry of tip portion
210 and grinding an oblique section, such as oblique section 300
and/or 320 into tip portion 210, reduces the wear of rotor blade
200 (shown in FIG. 3) when compared to baseline geometry 404
without the oblique section. Specifically, modifying tip portion
210 geometry reduces the rub ratio of blade 200. For example,
oblique section 300 within tip portion 210 alters the way in which
blade 200 contacts casing 106 during a rub event. Oblique section
300 lowers the contact force between rotor blade 200 and casing 106
thereby reducing vibration and deflection. By reducing the radial
and tangential loads induced into rotor blade 200, vibration is
reduced, thereby reducing radial elongation of rotor blade 200.
Additionally, modifying the geometry of tip portion 210 also
modifies the vibratory modes that contribute to radial elongation
within blade 200. Reducing radial elongation within rotor blade 200
decreases the amount of material loss due to rubbing against casing
106 and thus wear of rotor blade 200. In alternative embodiments,
modifying the geometry of tip portion 210 results in different rub
ratio values of blade 200 then illustrated in chart 400.
[0041] In the embodiments described above and referencing FIGS.
1-3, rotor blade 200 is shown and described as a compressor blade.
Within compressor section 104, each compressor stage may
incorporate rotor blades 200 that include different oblique
sections, such as oblique sections 300 and 320. For example, a
first compressor stage includes a plurality of rotor blades 200
with tip portion 210 having oblique section 300, while a second
compressor stage includes a plurality of rotor blades 200 with tip
portion 210 having oblique section 320. Moreover, in alternative
embodiments, tip portion 210 having an oblique section, such as
oblique section 300, is in any other blade within turbomachine 100,
such as, turbine section 112.
[0042] FIG. 5 is a schematic view of an alternative tip portion 500
for use with rotor blade 200 (shown in FIG. 2). In this alternative
embodiment, rotor blade 200 includes pressure sidewall 212 and an
opposing suction sidewall 214 which extend from root portion 208 to
tip portion 500. Additionally, tip portion 500 includes planar
section 222 that extends from pressure sidewall 212 to suction
sidewall 214 and substantially perpendicular thereto. Further, tip
portion 500 includes an oblique section 502 that convexly curves
from planar section 222 inward towards root portion 208 to leading
edge 216 forming a recess 504. Specifically, convex oblique section
502 extends a distance 506 along tip portion 500 from leading edge
216 approximately 30% of chord distance 304 of airfoil 202.
Additionally, convex oblique section 502 extends a depth 508 from
planar section 222. In alternative embodiments, convex oblique
section 502 extends for any other distance 506 and/or depth 508
that enables rotor blade 200 to function as described herein.
[0043] Additionally, or alternatively, in this alternative
embodiment, tip portion 500 includes oblique section 510 that
concavely curves from planar section 222 inward towards root
portion 208 to trailing edge 218 forming a recess 512.
Specifically, concave oblique section 510 extends a distance 514
along tip portion 500 from trailing edge 218 approximately 30% of
chord distance 304 of airfoil 202. Additionally, concave oblique
section 510 extends for depth 508 from planar section 222. In
alternative embodiments, concave oblique section 510 extends for
any other distance 514 and/or depth 508 that enables rotor blade
200 to function as described herein. Further, in alternative
embodiments, concave oblique section 510 is adjacent to leading
edge 216 and/or convex oblique section 502 is adjacent to trailing
edge 218.
[0044] Similar to tip portion 210 (shown in FIG. 3), tip portion
500 reduces the rub ratio of blade 200. Oblique section 502 and/or
510 lowers the contact force between rotor blade 200 and casing 106
(shown in FIG. 1) thereby reducing radial elongation. Reducing
radial elongation within rotor blade 200 decreases the amount of
material loss due to rubbing against casing 106 and thus wear of
rotor blade 200.
[0045] FIG. 6 is a schematic view of another alternative tip
portion 600 for use with rotor blade 200 (shown in FIG. 2). In this
alternative embodiment, rotor blade 200 includes pressure sidewall
212 and an opposing suction sidewall 214 which extend from root
portion 208 to tip portion 600. Additionally, tip portion 600
includes a first planar section 602 and a second planar section 604
that each extend from pressure sidewall 212 to suction sidewall 214
and substantially perpendicular thereto. Further, tip portion 600
includes an oblique section 606 forming a recess 608 between first
and second planar sections 602 and 604. Specifically, oblique
section 606 extends a distance 610 along tip portion 600 from first
planar section 602 to second planar section 604 at approximately
40% of chord distance 304 of airfoil 202 centering about mid-chord
line 217. Oblique section 606 has a first section 612 that extends
from first planar section 602 to mid-chord line 217 at a depth 614
such that first section 612 slopes from first planar section 602
towards root portion 208 in a direction towards trailing edge 218.
Oblique section 606 has a second section 616 that extends from
second planar section 604 to mid-chord line 217 such that second
section 616 slopes from second planar section 604 towards root
portion 208 in a direction towards leading edge 216. In this
alternative embodiment, oblique section 606 forms a V-shaped recess
608 about mid-chord line 217. In alternative embodiments, oblique
section 606 extends for any other distance 610 and/or depth 614
that enables rotor blade 200 to function as described herein.
Additionally, in alternative embodiments, oblique section 606 does
not center about mid-chord line 217.
[0046] Similar to tip portion 210 (shown in FIG. 3), tip portion
600 reduces the rub ratio of blade 200. Oblique section 606 lowers
the contact force between rotor blade 200 and casing 106 (shown in
FIG. 1) thereby reducing radial elongation. Reducing radial
elongation within rotor blade 200 decreases the amount of material
loss due to rubbing against casing 106 and thus wear of rotor blade
200.
[0047] FIG. 7 is a schematic view of a further alternative tip
portion 700 for use with rotor blade 200 (shown in FIG. 2). In this
alternative embodiment, rotor blade 200 includes pressure sidewall
212 and an opposing suction sidewall 214 which extend from root
portion 208 to tip portion 700. Additionally, tip portion 700
includes a first planar section 702 and a second planar section 704
that each extend from pressure sidewall 212 to suction sidewall 214
and substantially perpendicular thereto. Further, tip portion 700
includes an oblique section 706 forming a recess 708 between first
and second planar sections 702 and 704. Specifically, oblique
section 706 extends a distance 710 along tip portion 700 from first
planar section 702 to second planar section 704 at approximately
40% of chord distance 304 of airfoil 202 centering about mid-chord
line 217. Oblique section 706 has a first section 712 that extends
from first planar section 702 to mid-chord line 217 at a depth 714
such that first section 712 concavely slopes from first planar
section 602 towards root portion 208 in a direction towards
trailing edge 218. Oblique section 706 has a second section 716
that extends from second planar section 704 to mid-chord line 217
such that second section 716 convexly slopes from second planar
section 704 towards root portion 208 in a direction towards leading
edge 216. In this alternative embodiment, oblique section 706 forms
a U-shaped recess 708 about mid-chord line 217. In alternative
embodiments, oblique section 706 extends for any other distance 710
and/or depth 714 that enables rotor blade 200 to function as
described herein. Additionally, in alternative embodiments, oblique
section 706 does not center about mid-chord line 217.
[0048] Similar to tip portion 210 (shown in FIG. 3), tip portion
700 reduces the rub ratio of blade 200. Oblique section 706 lowers
the contact force between rotor blade 200 and casing 106 (shown in
FIG. 1) thereby reducing radial elongation. Reducing radial
elongation within rotor blade 200 decreases the amount of material
loss due to rubbing against casing 106 and thus wear of rotor blade
200.
[0049] The above described rotor blade tip geometries reduces wear
in a turbomachine. Specifically, a rotor blade includes an airfoil
having a suction sidewall coupled to a pressure sidewall at a
leading edge and a trailing edge. A tip portion extends between the
suction sidewall and the pressure sidewall and includes a planar
section and an oblique section. In some embodiments, the tip
portion includes a first oblique section and a second oblique
section. Modifying the rotor blade tip geometry by grinding the tip
portion and forming the oblique section reduces the rub ratio of
the rotor blade and, thereby, the wear of the rotor blade.
Specifically, the oblique section is sized such that a contact area
between the rotor blade and a surrounding casing is reduced,
thereby decreasing the radial and tangential loads induced into the
rotor blade during a rub event. Reducing the loads resulting from a
rub event decreases vibration and deflection of the rotor blade and
reduces material loss at the tip portion. Furthermore, modifying
the rotor blade tip geometry changes the vibratory modes of the
rotor blade such that radial elongation is decreased further
reducing material loss at the tip portion. Additionally, a
reduction in radial deflection allows the rotor blade to be
positioned closer to the surrounding casing. Accordingly,
decreasing the rub ratio of the rotor blade decreases wear and
material loss during a rub event, increases turbomachine
performance, and reduces maintenance costs.
[0050] An exemplary technical effect of the methods, systems, and
apparatus described herein includes at least one of the following:
(a) decreasing material loss of the rotor blade tip during a rub
event with a surrounding casing; (b) reducing wear of the rotor
blade; (b) decreasing a clearance gap between the rotor blade and
the casing; (c) reducing maintenance costs of turbomachines; and
(d) increasing turbomachine performance.
[0051] Exemplary embodiments of methods, systems, and apparatus for
reducing rotor blade tip wear are not limited to the specific
embodiments described herein, but rather, components of systems
and/or steps of the methods may be utilized independently and
separately from other components and/or steps described herein.
Further, the methods, systems, and apparatus may also be used in
combination with other systems requiring decreasing wear from a rub
event, and the associated methods are not limited to practice with
only the systems and methods described herein. Rather, the
exemplary embodiment can be implemented and utilized in connection
with many other applications, equipment, and systems that may
benefit from reducing wear on a blade tip.
[0052] Although specific features of various embodiments of the
disclosure may be shown in some drawings and not in others, this is
for convenience only. In accordance with the principles of the
disclosure, any feature of a drawing may be referenced and/or
claimed in combination with any feature of any other drawing.
[0053] This written description uses examples to disclose the
embodiments, including the best mode, and also to enable any person
skilled in the art to practice the embodiments, including making
and using any devices or systems and performing any incorporated
methods. The patentable scope of the disclosure 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 language of the claims.
* * * * *