U.S. patent application number 14/680382 was filed with the patent office on 2015-10-15 for ingestion blocking endwall feature.
The applicant listed for this patent is United Technologies Corporation. Invention is credited to Conway Chuong, Jeffrey J. Lienau, William Yeager.
Application Number | 20150292344 14/680382 |
Document ID | / |
Family ID | 53268594 |
Filed Date | 2015-10-15 |
United States Patent
Application |
20150292344 |
Kind Code |
A1 |
Chuong; Conway ; et
al. |
October 15, 2015 |
INGESTION BLOCKING ENDWALL FEATURE
Abstract
An endwall of a gas turbine engine section comprises a flow path
surface and a wall. The flow path surfaces defines an inner
diameter of a main flow path through the gas turbine engine
section, and terminates in an aft waterfall step. The wall extends
substantially radially inward from the waterfall step, and defines
a blockage feature that impedes airflow from the main flow path to
a secondary air cavity situated inward of the waterfall step.
Inventors: |
Chuong; Conway; (Manchester,
CT) ; Lienau; Jeffrey J.; (Wethersfield, CT) ;
Yeager; William; (Jupiter, FL) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation |
Hartford |
CT |
US |
|
|
Family ID: |
53268594 |
Appl. No.: |
14/680382 |
Filed: |
April 7, 2015 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
61978491 |
Apr 11, 2014 |
|
|
|
Current U.S.
Class: |
415/1 ;
415/176 |
Current CPC
Class: |
F05D 2270/17 20130101;
F05D 2240/80 20130101; Y02T 50/673 20130101; Y02T 50/60 20130101;
F01D 5/145 20130101; F01D 11/00 20130101; F01D 5/143 20130101; F05D
2250/11 20130101 |
International
Class: |
F01D 11/00 20060101
F01D011/00 |
Claims
1. An endwall of a gas turbine engine section, the endwall
comprising: a flow path surface defining an inner diameter of a
main flow path through the gas turbine engine section, and
terminating in an aft waterfall step; and a wall extending
substantially radially inward from the waterfall step, and defining
a blockage feature that impedes airflow from the main flow path to
a secondary air cavity situated inward of the waterfall step.
2. The endwall of claim 1, wherein the wall is formed via a
triangular cross-section widening of a region of the endwall near
the waterfall step.
3. The endwall of claim 2, wherein the triangular cross-section
widening comprises a substantially triangular cross-section portion
of additional material extending from an endwall into the secondary
air cavity.
4. The endwall of claim 3, wherein the substantially triangular
cross-section portion has a substantially right triangle
cross-section.
5. The endwall of claim 1, wherein the substantially radial wall is
radially thicker than an adjacent forward region of a second gas
turbine engine section immediately downstream of the gas turbine
engine section.
6. The endwall of claim 5, wherein the gas turbine section includes
a cooling air inlet into the secondary air cavity.
7. The endwall of claim 1, wherein the gas turbine engine section
is a vane section.
8. The endwall of claim 1, wherein the flow path surface is
substantially frustoconical, such that the flow path defined by the
flow path surface extends radially outward as it extends axially
aft.
9. A gas turbine engine comprising: a first gas turbine section
with a first platform defining an inner diameter of a main gas
turbine flow path in a first region; a second gas turbine section
situated immediately aft of the first gas turbine section, and
having a second platform defining an inner diameter of the main gas
turbine flow path in a second region; and an annular secondary air
cavity situated at an aft inner diameter of the first platform,
between the first and second gas turbine sections; an ingestion
reduction feature disposed to reduce airflow from the main gas
turbine flow path to the annular secondary air cavity, the
ingestion reduction feature comprising: a waterfall step at an aft
end of the first platform higher than an adjacent forward end of
the second platform; and a wall extending substantially radially
inward from the waterfall step, into the annular secondary air
cavity.
10. The gas turbine engine of claim 9, wherein the wall has radial
wall thickness greater than a radial thickness of an adjacent
forward tip of the second endwall.
11. The gas turbine engine of claim 10, wherein the wall extends
from an outer radial extent radially outward from an outer radial
extent of the adjacent forward tip, to an inner radial extent
radially inward from an inner radial extent of the adjacent forward
tip.
12. The gas turbine engine of claim 9, wherein the first turbine
section is a turbine exhaust case, and the second turbine section
is a power turbine.
13. The gas turbine engine of claim 9, wherein the first platform
and the second platform are both stationary endwalls.
14. The gas turbine engine of claim 9, further comprising a cooling
air inlet into the annular secondary air cavity.
15. The gas turbine engine of claim 9, wherein the wall is formed
via a triangular cross-section widening of the first platform at
the aft end of the first platform.
16. A method for reducing combustion gas ingestion into a secondary
air cavity of a gas turbine engine, the method comprising:
directing substantially laminar flow along a frustoconical flow
path surface terminating in a waterfall step disposed radially
outward of an adjacent tip of a downstream vane platform, such that
the combustion predominantly jumps from the frustoconical flow path
surface to the downstream vane platform without being ingested
therebetween; and impeding radially inward backflow between the
waterfall step and the downstream vane platform via a radial wall
extending radially inward from the waterfall step, and into the
secondary air cavity.
17. The method of claim 16, wherein the radial wall extends from a
radial location radially outward of an outer radial extent of the
adjacent tip, to a radial location radially inward of an inner
radial extent of the adjacent tip.
18. The method of claim 16, wherein a downstream projection of the
frustoconical surface extends radially outboard of an adjacent tip.
Description
CROSS-REFERENCE TO RELATED APPLICATION(S)
[0001] This application claims priority to U.S. Provisional
Application No. 61/978,491, filed on Apr. 11, 2014, and entitled
"INGESTION BLOCKING ENDWALL FEATURE", the disclosure of which is
incorporated by reference in its entirety.
BACKGROUND
[0002] The present invention relates generally to gas turbine
engine endwall features, and more specifically to endwall
configurations that reduce main gas path flow ingestion into
secondary, lower-temperature regions of the gas turbine engine.
[0003] Gas turbine engines operate according to a continuous-flow,
Brayton cycle. A compressor section pressurizes an ambient air
stream, fuel is added and the mixture is burned in a central
combustor section. The combustion products expand through a turbine
section where bladed rotors convert thermal energy from the
combustion products into mechanical energy for rotating one or more
centrally mounted shafts. The shafts, in turn, drive the forward
compressor section, thus continuing the cycle. Gas turbine engines
are compact and powerful power plants, making them suitable for
powering aircraft, heavy equipment, ships and electrical power
generators. In power generating applications, the combustion
products can also drive a separate power turbine attached to an
electrical generator.
[0004] Seals are required in many locations within a gas turbine
engine to regulate air flow to various portions of the engine. One
function of air seals in gas turbine engines is to limit ingestion
of hot "core" or "main" gas path airflow into secondary cavities
and passages. Secondary airflow system components are often
constructed to tolerate lower maximum temperatures than core gas
path components. Excessive ingestion of main gas path airflow can
result in undesirable heating of engine components, reducing part
lifetimes or necessitating that parts be constructed of expensive,
high temperature-capable materials.
SUMMARY
[0005] The present invention relates to an endwall of a gas turbine
engine section. The endwall comprises a flow path surface and a
wall. The flow path surface defines the inner diameter of a main
flow path through the gas turbine engine section, and terminates in
an aft waterfall step. The wall extends substantially radially
inward from the waterfall step, and defines a blockage feature that
impedes airflow from the main flow path to a secondary air cavity
situated inward of the waterfall step.
BRIEF DESCRIPTION OF THE DRAWINGS
[0006] FIG. 1 is a schematic view of a gas turbine engine.
[0007] FIG. 2 is a cross-sectional view of a turbine exhaust case
region of the gas turbine engine of FIG. 1.
[0008] FIG. 3 is a close-up cross-sectional view of an inner
diameter vane platform and cavity of the turbine exhaust case
region of FIG. 2.
[0009] FIG. 4 is a simplified cross-sectional view of the inner
diameter vane platform and cavity of FIG. 3, illustrating airflow
vectors.
[0010] FIG. 5 is a simplified cross-sectional view of the inner
diameter vane platform and cavity of FIG. 3, illustrating
temperatures.
[0011] While the above-identified figures set forth one or more
embodiments of the present disclosure, other embodiments are also
contemplated, as noted in the discussion. In all cases, this
disclosure presents the invention by way of representation and not
limitation. It should be understood that numerous other
modifications and embodiments can be devised by those skilled in
the art, which fall within the scope and spirit of the principles
of the invention. The figures may not be drawn to scale, and
applications and embodiments of the present invention may include
features and components not specifically shown in the drawings.
DETAILED DESCRIPTION
[0012] A flow blockage feature for a gas turbine engine can be
disposed at an aft inner diameter of a turbine exhaust case airflow
platform to reduce main airflow ingestion from a main engine gas
flow path into a radially inner secondary cavity. The flow blockage
feature includes a waterfall step and a substantially radial aft
platform wall that together reduce inward airflow between the
turbine exhaust case airflow platform and a downstream power
turbine airflow platform. The addition of the flow blockage feature
reduces the operating temperature of the secondary cavity, allowing
lower temperature-capable materials to be used.
[0013] An exemplary gas turbine engine 10 is circumferentially
disposed about a central, longitudinal axis or axial engine
centerline axis 12 as illustrated in FIG. 1. The engine 10 includes
in series order from front to rear, low and high pressure
compressor sections 16 and 18, a central combustor section 20 and
high and low pressure turbine sections 22 and 24. In some examples,
power turbine 26 is disposed aft of low pressure turbine 24.
Although illustrated with reference to an industrial gas turbine
engine, this application also extends to aero engines with a fan or
gear driven fan, and engines with more or fewer sections than
illustrated.
[0014] As is well known in the art of gas turbines, incoming
ambient air 30 becomes pressurized air 32 in the compressors 16 and
18. Fuel mixes with the pressurized air 32 in the combustor section
20, where it is burned to produce combustion gasses 34 that expand
as they flow through turbine sections 22, 24 and power turbine 26.
Turbine sections 22 and 24 drive high and low pressure rotor shafts
36 and 38 respectively, which rotate in response to the combustion
products and thus the attached compressor sections 18, 16. Power
turbine 26 can, for example, drive an electrical generator, pump,
gearbox, or other accessory (not shown). In the illustrated
embodiment, turbine section 24 meets power turbine 26 at joint
40.
[0015] It is understood that FIG. 1 provides a basic understanding
and overview of the various sections and the basic operation of an
industrial gas turbine engine. The illustrated embodiment is
provided merely by way of example and not limitation. Numerous
alternative configurations are possible, which may include
additional components not specifically shorn or omit certain
illustrated components. It will become apparent to those skilled in
the art that the present application is applicable to all types of
gas turbine engines, including those with aerospace
applications.
[0016] FIG. 2 shows a first module 42 and a second module 44 at
joint 40. First module 42 can, for example, be a turbine exhaust
case of turbine section 24, and second module 44 can, for example,
be a forward portion of power turbine 26. In the illustrated
embodiment, first and second modules 42 and 44 are vane sections of
gas turbine engine 10. In alternative embodiments, first and second
modules 42 and 44 can be other adjacent sections of gas turbine
engine 10, including mixed stationary and rotating components.
First module 42 and second module 44 are connected at joint 40 by
fasteners 45, which can for example be bolts, pins, or screws.
First module 42 is only partially illustrated in FIG. 2 and
includes a frame 46 and a fairing 48. Second module 44 includes
outer radial casing 47, a stator vane 50, vane platform 51, a rotor
blade 52, and a rotor disk 53. Frame 46 is a structural support
frame including outer frustoconical casing 54, inner
circumferential platform 56, and radial struts 58 extending
therebetween. Fairing 48 includes outer flow path platform 60, and
inner flow path platform 62, and strut liners 64. In the
illustrated embodiment, fairing 48 rides frame 46 and surrounds
struts 58 to define main engine gas flow path 68 through first
module 42. Outer flow path platform 60 and inner flow path platform
62 constitute annular endwalls of main engine flow path 68.
[0017] Inner flow path platform 62 terminates at aft inner diameter
(ID) platform feature 66, a flow blockage element that impedes
ingestion of combustion gasses 34 into secondary air cavity 74.
Secondary air cavity 74 is disposed between vane platform 51 of
second module 44, and seal support 67 and fairing 48 of first
module 42. Secondary air cavity 74 can, for example, receive
cooling air from inner plenum 70, an annular space extending from
inner circumferential platform 56 to inner flow path platform 62.
Aft ID platform feature 66 is described in greater detail below
with respect to FIGS. 3-5. Although aft ID platform feature 66 is
described with herein as a portion of fairing 48, aft ID platform
feature 66 can more generally be an aft portion of any similarly
situated endwall.
[0018] In the depicted embodiment, hot combustion gasses 34 flow
primarily through main engine gas flow path 68. Flow of combustion
gasses 34 into secondary air cavity 74, heats secondary air cavity
74. To reduce unwanted main flow gas path ingestion into secondary
air cavity 74, and thereby reduce the operating temperature of
secondary air cavity 74, inner flow path platform trailing edge is
disposed with aft ID feature 66.
[0019] FIG. 3 is a close-up cross-sectional view of region III of
gas turbine engine 10 from FIG. 2. Included in region III are
fairing 48, vane platform 51, inner flow path platform 62, aft ID
feature 66, seal support 67, inner plenum 70, and secondary air
cavity 74. Aft ID platform feature 66 of inner flow path platform
62 comprises frustoconical surface 100, waterfall step 102, and
radial wall 104. Adjacent tip 108 is a forwardmost tip of vane
platform 51 immediately downstream of radial wall 104, separated
from radial wall 104 by a small gap. Fairing 48 and seal support 67
meet at air seal 106, a partial sealing element that allows leakage
air to flow from inner plenum 70 in to secondary air cavity 74 via
leakage air inlet 106 and through exit 110. In the illustrated
embodiment, secondary air cavity 74 is further defined by finger
seal 112 and seal piece 114, although in alternative embodiments
secondary air cavity 74 can be bounded and defined by other
components.
[0020] Vane platform 51 and inner flow path platform 62 can be
constructed of high temperature-capable materials such as nickel or
cobalt-based superalloys, so as to withstand the high temperatures
of combustion gasses 34 passing through main engine gas flow path
68 (see FIG. 2, discussed above). Other elements surrounding
secondary air cavity 74 can be formed of less expensive, lower
temperature-capable materials. Leakage air from leakage air inlet
110 serves to reduce the overall temperature within secondary air
cavity 74, and aft ID platform feature 66 reduces ingestion of hot
combustion gasses 34 between inner flow path platform 62 and vane
platform 51, thereby obviating the need to use very high
temperature-capable materials in components surrounding secondary
air cavity 74.
[0021] Aft ID platform feature 66 comprises frustoconical surface
100, waterfall step 102, and radial wall 104. Frustoconical surface
100 extends to a maximum radius greater than the radius of vane
platform 51 at adjacent tip 108, terminating in waterfall step 102
(see FIG. 5). Waterfall step 102 thus discourages flow between
radial wall 104 and adjacent tip 108 by introducing a radial drop
from waterfall step 102 to adjacent tip 108. Projection line
P.sub.FR illustrates the slope of frustoconical surface 100
relative to adjacent tip 108. Substantially laminar flow along
frustoconical surface 100 tends to continue projection line
P.sub.FR, guided by waterfall step 102. Projection line P.sub.FR
extends radially outboard of adjacent tip 108 of vane platform 51.
Consequently, airflow through main engine gas flow path 68 tends to
pass from frustoconical surface 100 to vane platform 51,
effectively "jumping" the gap therebetween.
[0022] Radial wall 104 extends radially inward from waterfall step
102 into secondary air cavity 74, and is formed of additional
material thickening inner flow path platform 62 at its aft-most
extent. Radial wall 104 impedes forward flow of ingested combustion
gasses 34 from main engine gas flow path 68, thereby reducing
vortexing within secondary air cavity 74, and further decreasing
ingestion and corresponding heating of secondary air cavity 74. In
the illustrated embodiment, inner flow path platform 62 is of
substantially uniform thickness except at aft ID platform feature
66, where frustoconical surface 100 and radial step 104 together
define a substantially triangular cross-section to aft ID platform
feature 66. In one embodiment, radial wall 104 is formed via the
triangular cross-section widening of inner flow path platform 62 at
aft ID platform feature 66. This triangular cross-section can, for
example, be substantially a right triangle, as shown in FIG. 3.
More particularly, aft ID platform 66 comprises additional material
extending radially inward from the otherwise substantially flat
surface of inner flow path platform 62 into secondary air cavity 74
and terminating in radial wall 104.
[0023] FIG. 4 is a simplified cross-sectional view of the region
III (shown and described in greater detail with respect to FIG. 3,
above), with airflow direction vectors illustrating air flow
patterns in one embodiment of secondary air cavity 74. The airflow
patterns shown in FIG. 4 are intended for illustrative purposes
only; aft ID platform feature 66 can produce a wide range of
airflow patterns without departing from the spirit and scope of the
present invention.
[0024] In the illustrated embodiment, frustoconical surface 100,
waterfall step 102, and radial wall 104 of aft ID platform feature
66 together produce airflow patterns with low intensity vortices
V.sub.1, V.sub.2, V.sub.3, and V.sub.4. Although four vortices are
shown in FIG. 4, some embodiments of the present invention can
produce more or fewer vortices within secondary air cavity 74.
Generally, however, the airflow patterns produced by platform
feature 66 have several significant features. First, vortex airflow
within secondary air cavity 74 is characterized by low speed
relative to the flow of combustion gas 34 through main engine gas
flow path 68. Second, waterfall step 102 causes nearly laminar flow
along frustoconical surface 100 to predominantly continue to flow
downstream, "jumping" frustoconical surface 100 to vane platform 51
rather than turning radially inward towards secondary air cavity
(see projection line P.sub.FR, FIG. 3). Third, radial wall 104
blocks upstream airflow about vortex V.sub.0, further reducing
ingestion of this airflow. Aft ID platform feature 66 can, for
example, reduce ingestion of combustion gasses 34 from main engine
gas flow path 68 by approximately 40% relative to an uncontoured
inner flow path platform lacking aft ID platform feature 66. The
resulting airflow within secondary air cavity 74 is turbulent but
cool relative to combustion gasses 34. In some embodiments, air
from leakage air inlet 106 passes through leakage air exit 110 and
further cools secondary air cavity 74.
[0025] FIG. 5 is a simplified cross-sectional view of the region
III (shown and described in greater detail with respect to FIG. 3,
above), with isothermal lines showing temperature regions. FIG. 5
depicts only an illustrative example of temperature distributions
in and around secondary air cavity 74; actual temperature
distributions can vary across embodiments of the present invention
and as a function of operating conditions. In the illustrated
embodiment, region III has temperature zones labeled (in order of
increasing temperature) T.sub.1, T.sub.2, T.sub.3, T.sub.4, and
T.sub.5. Aft ID platform feature 66 increases the temperature
differential between combustion gasses 34 flowing through main
engine gas flow path 68 (e.g. region T.sub.5) and the secondary air
cavity 74 (predominantly region T.sub.1) by reducing ingestion of
combustion gasses 34 into secondary air cavity 74. As illustrated
in FIG. 5, aft ID platform feature 66 comprises frustoconical
surface 100, waterfall step 102, and radial wall 104. Frustoconical
surface 100 terminates at waterfall step 102, where aft ID platform
feature 66 turns inward in radial wall 104. Radial wall 104 has
wall height h.sub.w greater than an adjacent height h.sub.a of
adjacent tip 108. Moreover, radial wall 104 extends radially both
outward and inward of adjacent tip 108. Radial wall 104 thus
impedes ingestion of combustion gasses 34 by interrupting backflow
of gasses along adjacent tip 108 towards secondary air cavity 74.
Main (core) laminar air flow axially downstream along frustoconical
surface 100 predominantly "jumps" from aft ID platform feature 66
to vane platform 51, due to waterfall step 102. Although some such
gasses backflow (see vortex V.sub.0, FIG. 4), radial wall 104
redirects the bulk of such backflowing gasses back into main engine
gas flow path 68, producing a sharp temperature gradient between
main engine gas flow path 68 and secondary air cavity 74. Aft ID
platform feature 66 can, for example, reduce air temperatures in
secondary air cavity 74 by 50.degree. F., 75.degree. F., or more,
reducing the maximum operating temperatures of surrounding
components. By lowering the operating temperatures of components
around inner air cavity 74, aft ID platform feature 66 extends part
lifetimes, and allows components to be formed of less expensive,
lower temperature-capable materials.
Discussion of Possible Embodiments
[0026] The following are non-exclusive descriptions of possible
embodiments of the present invention.
[0027] An endwall of a gas turbine engine section, the endwall
comprising: a flow path surface defining an inner diameter of a
main flow path through the gas turbine engine section, and
terminating in an aft waterfall step; and a wall extending
substantially radially inward from the waterfall step, and defining
a blockage feature that impedes airflow from the main flow path to
a secondary air cavity situated inward of the waterfall step.
[0028] The endwall of the preceding paragraph can optionally
include, additionally and/or alternatively, any one or more of the
following features, configurations and/or additional
components:
[0029] A further embodiment of the foregoing endwall wherein the
wall is formed via a triangular cross-section widening of a region
of the endwall near the waterfall step.
[0030] A further embodiment of the foregoing endwall wherein the
triangular cross-section widening comprises a substantially
triangular cross-section portion of additional material extending
from an endwall into the secondary air cavity.
[0031] A further embodiment of the foregoing endwall wherein the
substantially triangular cross-section portion has a substantially
right triangle cross-section.
[0032] A further embodiment of the foregoing endwall wherein the
substantially radial wall is radially thicker than an adjacent
forward region of a second gas turbine engine section immediately
downstream of the gas turbine engine section.
[0033] A further embodiment of the foregoing endwall wherein the
gas turbine section includes a cooling air inlet into the secondary
air cavity.
[0034] A further embodiment of the foregoing endwall wherein the
gas turbine engine section is a vane section.
[0035] A further embodiment of the foregoing endwall wherein the
flow path surface is substantially frustoconical, such that the
flow path defined by the flow path surface extends radially outward
as it extends axially aft
[0036] A gas turbine engine comprising: a first gas turbine section
with a first platform defining an inner diameter of a main gas
turbine flow path in a first region; a second gas turbine section
situated immediately aft of the first gas turbine section, and
having a second platform defining an inner diameter of the main gas
turbine flow path in a second region; and an annular secondary air
cavity situated at an aft inner diameter of the first platform,
between the first and second gas turbine sections; an ingestion
reduction feature disposed to reduce airflow from the main gas
turbine flow path to the annular secondary air cavity, the
ingestion reduction feature comprising: a waterfall step at an aft
end of the first platform higher than an adjacent forward end of
the second platform; and a wall extending substantially radially
inward from the waterfall step, into the annular secondary air
cavity.
[0037] The gas turbine engine of the preceding paragraph can
optionally include, additionally and/or alternatively, any one or
more of the following features, configurations and/or additional
components:
[0038] A further embodiment of the foregoing gas turbine engine,
wherein the wall has radial wall thickness greater than a radial
thickness of an adjacent forward tip of the second endwall.
[0039] A further embodiment of the foregoing gas turbine engine,
wherein the wall extends from an outer radial extent radially
outward from an outer radial extent of the adjacent forward tip, to
an inner radial extent radially inward from an inner radial extent
of the adjacent forward tip.
[0040] A further embodiment of the foregoing gas turbine engine,
wherein the first turbine section is a turbine exhaust case, and
the second turbine section is a power turbine.
[0041] A further embodiment of the foregoing gas turbine engine,
wherein the first platform and the second platform are both
stationary endwalls.
[0042] A further embodiment of the foregoing gas turbine engine,
further comprising a cooling air inlet into the annular secondary
air cavity.
[0043] A further embodiment of the foregoing gas turbine engine,
wherein the wall is formed via a triangular cross-section widening
of the first platform at the aft end of the first platform.
[0044] A method for reducing combustion gas ingestion into a
secondary air cavity of a gas turbine engine, the method
comprising: directing substantially laminar flow along a
frustoconical flow path surface terminating in a waterfall step
disposed radially outward of an adjacent tip of a downstream vane
platform, such that the combustion predominantly jumps from the
frustoconical flow path surface to the downstream vane platform
without being ingested therebetween; and impeding radially inward
backflow between the waterfall step and the downstream vane
platform via a radial wall extending radially inward from the
waterfall step, and into the secondary air cavity.
[0045] The method of the preceding paragraph can optionally
include, additionally and/or alternatively, any one or more of the
following features, configurations and/or additional
components:
[0046] A further embodiment of the foregoing method, wherein the
radial wall extends from a radial location radially outward of an
outer radial extent of the adjacent tip, to a radial location
radially inward of an inner radial extent of the adjacent tip.
[0047] A further embodiment of the foregoing method, wherein a
downstream projection of the frustoconical surface extends radially
outboard of an adjacent tip.
[0048] Summation
[0049] Any relative terms or terms of degree used herein, such as
"substantially", "essentially", "generally", "approximately" and
the like, should be interpreted in accordance with and subject to
any applicable definitions or limits expressly stated herein. In
all instances, any relative terms or terms of degree used herein
should be interpreted to broadly encompass any relevant disclosed
embodiments as well as such ranges or variations as would be
understood by a person of ordinary skill in the art in view of the
entirety of the present disclosure, such as to encompass ordinary
manufacturing tolerance variations, incidental alignment
variations, alignment or shape variations induced by thermal,
rotational or vibrational operational conditions, and the like.
[0050] While the invention has been described with reference to an
exemplary embodiment(s), it will be understood by those skilled in
the art that various changes may be made and equivalents may be
substituted for elements thereof without departing from the scope
of the invention. In addition, many modifications may be made to
adapt a particular situation or material to the teachings of the
invention without departing from the essential scope thereof.
Therefore, it is intended that the invention not be limited to the
particular embodiment(s) disclosed, but that the invention will
include all embodiments falling within the scope of the appended
claims.
* * * * *