U.S. patent application number 13/664031 was filed with the patent office on 2014-06-05 for endwall controuring.
This patent application is currently assigned to United Technologies Corporation. The applicant listed for this patent is United Technologies Corporation. Invention is credited to Andrew S. Aggarwala, Eunice Allen-Bradley, Eric A. Grover, Timothy Charles Nash.
Application Number | 20140154068 13/664031 |
Document ID | / |
Family ID | 50685292 |
Filed Date | 2014-06-05 |
United States Patent
Application |
20140154068 |
Kind Code |
A1 |
Aggarwala; Andrew S. ; et
al. |
June 5, 2014 |
Endwall Controuring
Abstract
An airfoil array is disclosed. The airfoil array may include an
endwall, and a plurality of airfoils radially projecting from the
endwall. Each airfoil may have a first side and an opposite second
side extending axially in chord between a leading edge and a
trailing edge. The airfoils may be circumferentially spaced apart
on the endwall thereby defining a plurality of flow passages
between adjacent airfoils. The airfoil array may further include a
convex profiled region extending from the endwall adjacent to the
first side of at least one of said plurality of airfoils near the
leading edge of the at least one of said plurality of airfoils, and
a concave profiled region in the endwall and extending across said
at least one of said plurality of flow passages.
Inventors: |
Aggarwala; Andrew S.;
(Vernon, CT) ; Nash; Timothy Charles; (East
Hartford, CT) ; Allen-Bradley; Eunice; (Vernon,
CT) ; Grover; Eric A.; (Tolland, CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation; |
|
|
US |
|
|
Assignee: |
United Technologies
Corporation
Hartford
CT
|
Family ID: |
50685292 |
Appl. No.: |
13/664031 |
Filed: |
October 30, 2012 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
61706983 |
Sep 28, 2012 |
|
|
|
Current U.S.
Class: |
415/208.1 ;
416/243 |
Current CPC
Class: |
Y02T 50/673 20130101;
F01D 5/145 20130101; F04D 29/324 20130101; F04D 29/681 20130101;
F04D 29/542 20130101; F01D 9/02 20130101; F01D 5/143 20130101; Y02T
50/60 20130101 |
Class at
Publication: |
415/208.1 ;
416/243 |
International
Class: |
F01D 5/14 20060101
F01D005/14; F01D 9/02 20060101 F01D009/02 |
Claims
1. An airfoil array, comprising: an endwall; a plurality of
airfoils radially projecting from the endwall, each airfoil having
a first side and an opposite second side extending axially in chord
between a leading edge and a trailing edge, the airfoils
circumferentially spaced apart on the endwall thereby defining a
plurality of flow passages between adjacent airfoils; a convex
profiled region extending from the endwall adjacent the first side
of at least one of said plurality of airfoils near the leading edge
of the at least one of said plurality of airfoils; and a concave
profiled region in the endwall and extending across said at least
one of said plurality of flow passages.
2. The airfoil array of claim 1, wherein a local maximum in radial
extent of the convex profiled region is positioned between the
leading edge and mid-chord of the airfoil.
3. The airfoil array of claim 1, wherein each airfoil has an axial
chord, and a local maximum in radial extent of the convex profiled
region is disposed between about 0% to about 50% of the axial
chord.
4. The airfoil array of claim 1, wherein each flow passage has a
passage width, and a local maximum in radial extent of the convex
profiled region is disposed between about 0% to about 50% of the
passage width.
5. The airfoil array of claim 1, wherein a local minimum in radial
extent of the concave profiled region is positioned between the
leading edge and the trailing edge of the airfoil.
6. The airfoil array of claim 1, wherein the concave profiled
region is elongated across the flow passage from the first side of
at least one of said plurality of airfoils to the second side of an
adjacent airfoil.
7. The airfoil array of claim 6, wherein the concave profiled
region is adjacent to the first side of the airfoil near
mid-chord.
8. The airfoil array of claim 6, wherein the concave profiled
region is adjacent to the second side of the adjacent airfoil along
a significant length of the second side.
9. The airfoil array of claim 6, wherein the concave profiled
region adjacent to the second side of the adjacent airfoil extends
along a majority of the second side of the airfoil.
10. The airfoil array of claim 6, wherein an extent of the concave
profiled region adjacent the first side of the airfoil is less than
an extent of the concave profiled region adjacent the second side
of the adjacent airfoil.
11. The airfoil array of claim 1, wherein each airfoil has an axial
chord and each flow passage has a passage width, and a local
minimum in radial extent of the concave profiled region is disposed
between about 0% to about 100% of the axial chord and between about
0% to about 100% of the passage width.
12. The airfoil array of claim 1, wherein the first side is a
pressure side of an airfoil, and the opposite second side is a
suction side of an airfoil.
13. A gas turbine engine, comprising: a compressor section; a
combustor section downstream of the compressor section; and a
turbine section downstream of the combustor section, one of the
compressor section and the turbine section having at least one
airfoil array including a plurality of airfoils circumferentially
spaced apart and projecting radially from an endwall, the airfoils
establishing a plurality of flow passages, each airfoil having a
first side, an opposite second side, a leading edge, and a trailing
edge, the endwall having a convex profiled surface near the leading
edge of at least one of said plurality of airfoils, and a concave
profiled surface elongated across at least one of said plurality of
flow passages from the first side of the at least one of said
plurality of airfoils to the second side of an adjacent
airfoil.
14. The gas turbine engine of claim 13, wherein the convex profiled
surface is adjacent to the first side of the at least one of said
plurality of airfoils.
15. The gas turbine engine of claim 13, wherein each airfoil has an
axial chord, and the concave profiled surface adjacent to the first
side of the at least one of said plurality of airfoils extends
between about 30% to about 80% of the axial chord.
16. The gas turbine engine of claim 13, wherein each airfoil has an
axial chord, and the concave profiled surface adjacent to the
second side of the adjacent airfoil extends between about 0% to
about 100% of the axial chord.
17. The gas turbine engine of claim 13, wherein each airfoil has an
axial chord, and a local minimum in radial extent of the concave
profiled surface adjacent to the second side of the adjacent
airfoil is positioned between about 0% to about 100% of the axial
chord.
18. The gas turbine engine of claim 13, wherein each flow passage
has a passage width, and a local minimum in radial extent of the
concave profiled surface is disposed between about 30% to about 80%
of the passage width.
19. The gas turbine engine of claim 13, wherein a local maximum in
radial extent of the convex profiled surface is disposed between
about 0% to about 50% of the passage width, and a local minimum in
radial extent of the concave profiled surface is disposed between
about 50% to about 100% of the passage width.
20. The gas turbine engine of claim 13, wherein the first side is a
pressure side of an airfoil and the opposite second side is a
suction side of an airfoil.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] The present disclosure claims the benefit of priority from
U.S. Provisional Patent Application Ser. No. 61/706,983, filed on
Sep. 28, 2012, which is incorporated herein by reference.
FIELD OF THE DISCLOSURE
[0002] The present disclosure relates generally to airfoil arrays
utilized in gas turbine engines and, more particularly, to endwall
contouring.
BACKGROUND OF THE DISCLOSURE
[0003] Gas turbine engines typically include a compressor section,
a combustor section, and a turbine section, with an annular flow
path extending axially through each. Initially, air flows through
the compressor where it is compressed or pressurized. The combustor
then mixes and ignites the compressed air with fuel, generating hot
combustion gases. These hot combustion gases are then directed by
the combustor to the turbine where power is extracted from the hot
gases by causing blades of the turbine to rotate.
[0004] Some sections of the engine include airfoil arrays. Air
within the engine moves through fluid flow passages in the arrays.
The fluid flow passages are established by adjacent airfoils
projecting from laterally extending endwalls. Near the endwalls,
the fluid flow is dominated by a flow phenomenon known as a
horseshoe vortex, which forms as a result of the endwall boundary
layer separating from the endwall as the gas approaches the leading
edges of the airfoils. The separated gas reorganizes into the
horseshoe vortex. There is a high loss of efficiency associated
with the vortex, and this loss is referred to as "secondary" or
endwall loss. Accordingly, there exists a need for a way to
mitigate or reduce endwall losses.
SUMMARY OF THE DISCLOSURE
[0005] According to one embodiment of the present disclosure, an
airfoil array is disclosed. The airfoil array may comprise an
endwall, and a plurality of airfoils radially projecting from the
endwall. Each airfoil may have a first side and an opposite second
side extending axially in chord between a leading edge and a
trailing edge. The airfoils may be circumferentially spaced apart
on the endwall thereby defining a plurality of flow passages
between adjacent airfoils. The airfoil array may further comprise a
convex profiled region extending from the endwall adjacent to the
first side of at least one of said plurality of airfoils near the
leading edge of the at least one of said plurality of airfoils, and
a concave profiled region in the endwall and extending across said
at least one of said plurality of flow passages.
[0006] In a refinement, a local maximum in radial extent of the
convex profiled region may be positioned between the leading edge
and mid-chord of the airfoil.
[0007] In another refinement, each airfoil may have an axial chord,
and a local maximum in radial extent of the convex profiled region
may be disposed between about 0% to about 50% of the axial
chord.
[0008] In another refinement, each flow passage may have a passage
width, and a local maximum in radial extent of the convex profiled
region may be disposed between about 0% to about 50% of the passage
width.
[0009] In another refinement, a local minimum in radial extent of
the concave profiled region may be positioned between the leading
edge and the trailing edge of the airfoil.
[0010] In yet another refinement, the concave profiled region may
be elongated across the flow passage from the first side of at
least one of said plurality of airfoils to the second side of an
adjacent airfoil.
[0011] In a related refinement, the concave profiled region may be
adjacent to the first side of the airfoil near mid-chord.
[0012] In another related refinement, the concave profiled region
may be adjacent to the second side of the adjacent airfoil along a
significant length of the second side.
[0013] In another related refinement, the concave profiled region
adjacent to the second side of the adjacent airfoil may extend
along a majority of the second side of the airfoil.
[0014] In yet another related refinement, an extent of the concave
profiled region adjacent the first side of the airfoil may be less
than an extent of the concave profiled region adjacent the second
side of the adjacent airfoil.
[0015] In another refinement, each airfoil may have an axial chord
and each flow passage may have a passage width. A local minimum in
radial extent of the concave profiled region may be disposed
between about 0% to about 100% of the axial chord and between about
0% to about 100% of the passage width.
[0016] In yet another refinement, the first side may be a pressure
side of an airfoil, and the opposite second side may be a suction
side of an airfoil.
[0017] According to another embodiment, a gas turbine engine is
disclosed. The gas turbine engine may comprise a compressor
section, a combustor section downstream of the compressor section,
and a turbine section downstream of the combustor section. One of
the compressor section and the turbine section may have at least
one airfoil array including a plurality of airfoils
circumferentially spaced apart and projecting radially from an
endwall, the airfoils establishing a plurality of flow passages.
Each airfoil may have a first side, an opposite second side, a
leading edge, and a trailing edge. The endwall may have a convex
profiled surface near the leading edge of at least one of said
plurality of airfoils, and a concave profiled surface elongated
across at least one of said plurality of flow passages from the
first side of the at least one of said plurality of airfoils to the
second side of an adjacent airfoil.
[0018] In a refinement, the convex profiled surface may be adjacent
to the first side of each airfoil.
[0019] In another refinement, each airfoil may have an axial chord,
and the concave profiled surface adjacent to the first side of each
airfoil may extend between about 30% to about 80% of the axial
chord.
[0020] In another refinement, each airfoil may have an axial chord,
and the concave profiled surface adjacent to the second side of the
adjacent airfoil may extend between about 0% to about 100% of the
axial chord.
[0021] In another refinement, each airfoil may have an axial chord,
and a local minimum in radial extent of the concave profiled
surface adjacent to the second side of each airfoil may be
positioned between about 0% to about 100% of the axial chord.
[0022] In another refinement, each flow passage may have a passage
width, and a local minimum in radial extent of the concave profiled
surface may be disposed between about 30% to about 80% of the
passage width.
[0023] In another refinement, a local maximum in radial extent of
the convex profiled surface may be disposed between about 0% to
about 50% of the passage width, and a local minimum in radial
extent of the concave profiled surface may be disposed between
about 50% to about 100% of the passage width.
[0024] In yet another refinement, the first side may be a pressure
side of an airfoil and the opposite second side may be a suction
side of an airfoil.
[0025] These and other aspects and features of the disclosure will
become more readily apparent upon reading the following detailed
description when taken in conjunction with the accompanying
drawings.
BRIEF DESCRIPTION OF THE DRAWINGS
[0026] FIG. 1 is a partial sectional view of a gas turbine engine
according to one embodiment of the present disclosure;
[0027] FIG. 2 is a perspective view of an airfoil array within the
gas turbine engine of FIG. 1; and
[0028] FIG. 3 is a plan view with topographic contours showing a
portion of the airfoil array of FIG. 2.
[0029] While the present disclosure is susceptible to various
modifications and alternative constructions, certain illustrative
embodiments thereof, will be shown and described below in detail.
It should be understood, however, that there is no intention to be
limited to the specific embodiments disclosed, but on the contrary,
the intention is to cover all modifications, alternative
constructions, and equivalents along within the spirit and scope of
the present disclosure.
DETAILED DESCRIPTION
[0030] Referring now to the drawings, and with specific reference
to FIG. 1, in accordance with the teachings of the disclosure, an
exemplary gas turbine engine 10 is shown. The gas turbine engine 10
may generally comprise a compressor section 12 where air is
pressurized, a combustor section 14 downstream of the compressor
section 12 which mixes and ignites the compressed air with fuel,
thereby generating hot combustion gases, a turbine section 16
downstream of the combustor section 14 for extracting power from
the hot combustion gases, and an annular flow path 18 extending
axially through each.
[0031] The turbine section 16 or the compressor section 12 may
include at least one airfoil array 20. As shown best in FIG. 2, the
airfoil array 20 may comprise a plurality of airfoils 22 projecting
radially from an endwall 24. For example, the airfoils 22 may be
provided as a stage of rotor blades or stator vanes in the
compressor section 12 or the turbine section 16 of the gas turbine
engine 10. The endwall 24 may be either an inner diameter (ID)
endwall or an outer diameter (OD) endwall or both. The airfoils 22
may be circumferentially spaced apart on the endwall 24 and
arranged about the engine centerline X (FIG. 1), thereby defining a
plurality of fluid flow passages 26 between adjacent airfoils 22
with the endwall 24. Each airfoil 22 may have a first side 28 and
an opposite second side 30 extending axially in chord between a
leading edge 32 and a trailing edge 34. Fluid flow, such as
airflow, moves toward the flow passage 26 from a position forward
of the leading edge 32 of the airfoils 22 as the engine 10
operates.
[0032] The endwall 24 may have a plurality of convex profiled
regions 36 and a plurality of concave profiled regions 38
configured to direct flow through each of the flow passages 26
within the airfoil array 20. Illustrated in FIG. 3 with topographic
contour lines, near the leading edge 32 of at least one of said
plurality of airfoils 22 within the airfoil array 20, the convex
profiled region 36 may be located adjacent to the first side 28 of
each airfoil 22. The convex profiled region 36 may extend radially
inward toward the annular flow path 18.
[0033] Each airfoil 22 may have a chord 42, which is defined as a
line from the leading edge 32 to the trailing edge 34, and an axial
chord 44, which is a projection of the chord 42 onto a plane
containing the engine centerline X. Relevant distances may be
expressed as a percentage of the length of the axial chord 44, as
shown in the percentage scale at the bottom of FIG. 3. Each fluid
flow passage 26 may have a passage width W measured from the first
side 28 of each airfoil 22 to the second side 30 of a neighboring
airfoil 46. The passage width W may typically vary from a passage
inlet 48 to a passage outlet 50 so that the passage width may be
locally different at different chordwise locations. Relevant
distances may be expressed as a fraction or percentage of the
length of the passage width W, with 0% referenced at the first side
28 of each airfoil 22 and 100% referenced at the second side 30 of
the neighboring airfoil 46.
[0034] From the surface 40 adjacent the convex profiled region 36,
the convex region 36 may gradually increase in radial height, or
move radially inward toward the annular flow path 18, to a local
maximum in radial extent 52. It will be understood that the convex
profiled region 36 may extend further than the illustrated contour
lines. The convex profiled region 34 may have the local maximum in
radial extent 52 positioned between the leading edge 32 and
mid-chord of the airfoil 22. The local maximum in radial extent 52
of the convex profiled region 34 may be disposed within an
inclusive axial range of about 0% to about 50% of the axial chord
and may be disposed within an inclusive lateral range of about 0%
to about 50% of the passage width W.
[0035] The concave profiled region 38 may be elongated across at
least one of said plurality of fluid flow passages 26. Relative to
a surface 40 adjacent the concave profiled region 38, the concave
region 38 may extend radially outward away from the annular flow
path 18. From the surface 40 adjacent the concave profiled region
38, the concave region 38 may gradually decrease in radial height,
or move outward radially away from the annular flow path 18, to a
local minimum in radial extent 54. The concave region 38 of the
endwall 24 may extend across the fluid flow passage 26 from the
first side 28 of each airfoil 22 to the second side 30 of the
neighboring airfoil 46. It will be understood that the concave
profiled region 38 may extend further than the illustrated contour
lines.
[0036] The concave profiled region 38 may be adjacent to the first
side 28 of the airfoil 22 near mid-chord and may be predominantly
adjacent to the second side 30 of the neighboring airfoil 46 along
a significant length of the second side 30, about a majority of the
second side 30 of the airfoil 46. The extent of the concave region
38 adjacent the first side 28 of the airfoil 22 may be less than
the extent of the concave region 38 adjacent the second side 30 of
the neighboring airfoil 46. For example, adjacent the first side 28
of the airfoil 22, the concave region 38 may extend between an
inclusive range of about 30% to about 80% of the axial chord.
Adjacent the second side 30 of the neighboring airfoil 46, the
concave region 38 may extend between an inclusive range of about 0%
to about 100% of the axial chord. The local minimum in radial
extent 54 of the concave profiled region 38 may be positioned
between the leading edge 32 and the trailing edge 34 of the
airfoils 22, 36, or within the inclusive range of about 0% to about
100% of the axial chord. The local minimum in radial extent 54 may
also be disposed within an inclusive lateral range of about 0% to
about 100% of the passage width, for example, between about 30% to
about 80% of the passage width, or for example, between about 50%
to about 100% of the passage width.
[0037] It will be understood that the endwall 24 contouring
described herein may be applied to any type of airfoil array 20
without departing from the scope of the invention. According to an
exemplary embodiment, the contoured endwall 24 may be applied to an
airfoil array 20 with airfoils 22 having a camber or turning
airfoil, as shown in FIG. 2. For example, the first side 28 of each
airfoil 22 may be a pressure side, the second side 30 of each
airfoil 22 may be a suction side, the convex profiled region 36 may
be adjacent the pressure side of the airfoil 22, and the concave
profiled region 38 may be elongated across each fluid flow passage
26 from the pressure side of each airfoil 22 to the suction side 30
of the neighboring airfoil. According to another exemplary
embodiment, the contoured endwall 24 may be applied to an airfoil
array 20 with airfoils 22 having no camber, as shown in FIG. 3.
Such airfoils 22 may be provided, for example, as compressor or
turbine blades or vanes, or middle turbine frames in a gas turbine
engine.
INDUSTRIAL APPLICABILITY
[0038] From the foregoing, it can be seen that the teachings of
this disclosure can find industrial application in any number of
different situations, including but not limited to, gas turbine
engines. Such engines may be used, for example, on aircraft for
generating thrust, or in land, marine, or aircraft applications for
generating power.
[0039] The disclosure described provides a way to mitigate or
reduce endwall losses in an airfoil array. By positioning a convex
profiled region of the endwall adjacent the first side of the
airfoils near the leading edge and a concave profiled region of the
endwall elongated across the flow passages, the present invention
influences flow through the flow passages, thereby reducing endwall
losses due to the horseshoe vortex. Furthermore, the contoured
endwall described herein results in an improved aerodynamic
performance of the airfoil arrays. Such contouring may minimize
aerodynamic losses through blade or vane passages of a gas turbine
engine, for example, those of a second stage high pressure turbine
vane. In so doing, this may decrease heat, friction and pressure
losses, while improving engine efficiency and life of the blade or
vane.
[0040] While the foregoing detailed description has been given and
provided with respect to certain specific embodiments, it is to be
understood that the scope of the disclosure should not be limited
to such embodiments, but that the same are provided simply for
enablement and best mode purposes. The breadth and spirit of the
present disclosure is broader than the embodiments specifically
disclosed and encompassed within the claims appended hereto.
* * * * *