Note : Les descriptions sont présentées dans la langue officielle dans laquelle elles ont été soumises.
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PCT11JS2011/067393
GAS TURBINE ENGINE AND VARIABLE CAMBER VANE SYSTEM
Government Rights
The present application was made with the United States government support
under Contract No. FA8650-07-C-2803, awarded by the United States Air Force.
The
United States government may have certain rights in the present application.
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Field of the Invention
The present invention relates to gas turbine engines, and more particularly,
to
gas turbine engines with variable camber vane systems.
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Background
Gas turbine engines with variable camber vane systems remain an area of
interest. Some existing systems have various shortcomings, drawbacks, and
disadvantages relative to certain applications. Accordingly, there remains a
need for
further contributions in this area of technology.
3
,
Summary
One embodiment of the present invention is a unique variable camber vane
system for a gas turbine
engine. Another embodiment is a unique gas turbine engine.
Other embodiments include apparatuses, systems, devices, hardware, methods,
and combinations for
gas turbine engines and variable camber vane systems. Further embodiments,
forms, features, aspects,
benefits, and advantages of the present application will become apparent from
the description and
figures provided herewith.
In accordance with an aspect of the present disclosure there is provided a
variable camber vane system
for a gas turbine engine, comprising: a first airfoil portion having a first
tip portion, a first root portion,
a face extending at least partially between the first tip portion and the
first root portion, and a groove
in the face extending at least partially between the first tip portion and the
first root portion, wherein
the groove has a groove width; a second airfoil portion arranged to rotate
with respect to the first airfoil
portion about a pivot axis, wherein the second airfoil portion includes a
second tip portion; a second
root portion; and a crown extending at least partially between the second tip
portion and the second
root portion, wherein the crown includes a crown radius centered about the
pivot axis and positioned
opposite the groove; and a seal strip having a seal width greater than the
groove width and a concave
rubbing surface, the rubbing surface having a radius complementary to and
opposite the crown radius,
wherein the seal strip is at least partially disposed in the groove with an
interference fit; and wherein
the seal strip is arranged to seal against fluid flow between the first
airfoil portion and the second airfoil
portion, wherein the seal strip is a rigid structure formed to maintain the
concave rubbing surface in
response to rotation of the second airfoil portion about the pivot axis.
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Brief Description of the Drawings
The description herein makes reference to the accompanying drawings wherein
like reference numerals refer to like parts throughout the several views, and
wherein:
FIG. 1 schematically depicts some aspects of a non-limiting example of a gas
turbine engine in accordance with an embodiment of the present invention.
FIG. 2 schematically depicts some aspects of a non-limiting example of a fan
system for a gas turbine engine in accordance with an embodiment of the
present
invention.
FIG. 3 depicts some aspects of a non-limiting example of a variable camber
guide vane system in accordance with an embodiment of the present invention.
FIG. 4 depicts some aspects of the variable camber guide vane system of FIG.
3.
FIG. 5 depicts some aspects of a non-limiting example of a seal strip in
accordance with an embodiment of the present invention.
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Detailed Description
For purposes of promoting an understanding of the principles of the invention,
reference will now be made to the embodiments illustrated in the drawings, and
specific
language will be used to describe the same. It will nonetheless be understood
that no
limitation of the scope of the invention is intended by the illustration and
description
of certain embodiments of the invention. In addition, any alterations and/or
modifications of the illustrated and/or described embodiment(s) are
contemplated as
being within the scope of the present invention. Further, any other
applications of the
principles of the invention, as illustrated and/or described herein, as would
normally
occur to one skilled in the art to which the invention pertains, are
contemplated as being
within the scope of the present invention.
Referring to the drawings, and in particular FIG. 1, a non-limiting example of
a
gas turbine engine 10 in accordance with an embodiment of the present
invention is
depicted. In one form, gas turbine engine 10 is an aircraft propulsion power
plant. In
other embodiments, gas turbine engine 10 may be a land-based or marine engine.
In
one form, gas turbine engine 10 is a multi-spool turbofan engine. In other
embodiments, gas turbine engine 10 may be a single or multi-spool turbofan,
turboshaft,
turbojet, turboprop gas turbine or combined cycle engine.
Gas turbine engine 10 includes a fan system 12, a compressor system 14, a
diffuser 16, a combustion system 18 and a turbine system 20. Compressor system
14
is in fluid communication with fan system 12. Diffuser 16 is in fluid
communication with
compressor system 14. Combustion system 18 is fluidly disposed between
compressor
system 14 and turbine system 20. Fan system 12 includes a fan rotor system 22.
In
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various embodiments, fan rotor system 22 includes one or more rotors (not
shown) that
are powered by turbine system 20. Compressor system 14 includes a compressor
rotor
system 24. In various embodiments, compressor rotor system 24 includes one or
more
rotors (not shown) that are powered by turbine system 20. Turbine system 20
includes
a turbine rotor system 26. In various embodiments, turbine rotor system 26
includes
one or more rotors (not shown) operative to drive fan rotor system 22 and
compressor
rotor system 24. Turbine rotor system 26 is driving coupled to compressor
rotor system
24 and fan rotor system 22 via a shafting system 28. In various embodiments,
shafting
system 28 includes a plurality of shafts that may rotate at the same or
different speeds
and directions. In some embodiments, only a single shaft may be employed.
During the operation of gas turbine engine 10, air is drawn into the inlet of
fan 12
and pressurized by fan 12. Some of the air pressurized by fan 12 is directed
into
compressor system 14, and the balance is directed into a bypass duct (not
shown).
Compressor system 14 further pressurizes the air received from fan 12, which
is then
discharged into diffuser 16. Diffuser 16 reduces the velocity of the
pressurized air, and
directs the diffused airflow into combustion system 18. Fuel is mixed with the
pressurized air in combustion system 18, which is then combusted. In one form,
combustion system 18 includes a combustion liner (not shown) that contains a
continuous combustion process. In other embodiments, combustion system 18 may
take other forms, and may be, for example, a wave rotor combustion system, a
rotary
valve combustion system, or a slinger combustion system, and may employ
deflagration
and/or detonation combustion processes. The hot gases exiting combustor 18 are
directed into turbine system 20, which extracts energy in the form of
mechanical shaft
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power to drive fan system 12 and compressor system 14 via shafting system 28.
The
hot gases exiting turbine system 20 are directed into a nozzle (not shown),
and provide
a component of the thrust output by gas turbine engine 10.
Referring to FIG. 2, a non-limiting example of some aspects of fan system 12
in
accordance with an embodiment of the present invention is schematically
depicted. Fan
system 12 includes a variable guide vane system 40 having a variable inlet
guide vane
stage 42 and a variable outlet guide vane stage 44 disposed on either side of
a rotating
fan stage 46. Variable inlet guide vane stage 42 is operative to guide air
into rotating
fan stage 46, and to selectively vary the incidence angle of the air flow
entering rotating
fan stage 46. Variable outlet guide vane stage 44 is operative to guide air
exiting
rotating fan stage 46, and to selectively vary the incidence angle of the air
flow exiting
rotating fan stage 46. Variable inlet guide vane stage 42 and variable outlet
guide vane
stage 44 are actuated by an actuation system (not shown). Although described
herein
as with respect to fan system 12, it will be understood that variable guide
vane system
40 may also or alternatively be employed as part of compressor system 14. In
addition,
although variable guide vane system 40 includes both variable inlet and outlet
guide
vane stages, other embodiments may include only a variable inlet guide vane
stage or a
variable outlet guide vane stage.
Referring to FIGS. 3-5, a non-limiting example of some aspects of variable
inlet
guide vane stage 42 in accordance with an embodiment of the present invention
is
illustrated. It will be understood that some embodiments of variable outlet
guide vane
stage 44 may be similar to variable inlet guide vane stage 42, and hence, the
following
description of variable inlet guide vane stage 42 is also applicable to
aspects of some
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embodiments of variable outlet guide vane stage 44. Variable inlet guide vane
stage 42
includes an outer band 50, an inner band 52 and plurality of vanes 54. Outer
band 50
defines an outer flowpath wall of variable inlet guide vane stage 42. Inner
band 52
defines an inner flowpath wall of variable inlet guide vane stage 42. Vanes 54
are
airfoils that extend between outer band 50 and inner band 52, and are spaced
apart
circumferentially. In one form, vanes 54 extend in the radial direction
between outer
band 50 and inner band 52. In other embodiments, vanes 54 may extend between
outer band 50 and inner band 52 at other angles.
Each vane 54 includes an airfoil portion 56 and an airfoil portion 58. Airfoil
portion 56 extends between a tip portion 60 and a root portion 62. In one
form, airfoil
portion 56 includes the trailing edge 64 of vane 54. In other embodiments,
airfoil portion
56 may be formed with a leading edge of vane 54 instead of trailing edge 64,
e.g., for
use in variable outlet guide vane 44. Airfoil portion 58 extends between a tip
portion 66
and a root portion 68. In one form, airfoil portion 58 includes the leading
edge 70 of
vane 54. In other embodiments, airfoil portion 58 may be formed with a
trailing edge
instead of leading edge 70, e.g., for use in variable outlet guide vane 44. In
one form,
airfoil portion 56 is fixed, i.e., stationary. In other embodiments, airfoil
portion 56 may
be movable, e.g., pivotable about an axis so as to be able to vary the angle
of the
trailing edge of vane 54. In one form, airfoil portion 58 is variable, being
configured to
pivot about a pivot axis 72 with respect to airfoil portion 56, to provide a
variable camber
for vane 54. In other embodiments, airfoil portion 58 may be fixed. In one
form, airfoil
portion 58 is coupled to an actuation system (not shown) that is operative to
selectively
position airfoil portion 58 at a desired incidence angle. In other
embodiments, airfoil
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portion 56 may also or alternatively be coupled to an actuation system (not
shown) that
is operative to selectively position airfoil portion 56 at a desired incidence
angle.
Extending from airfoil portion 58 are pivot shafts 74 and 76, which establish
pivot
axis 72. Outer band 50 includes a plurality of spaced apart openings 78. Inner
band 52
includes a plurality of spaced apart openings 80. Openings 78 and 80 are
operative to
receive pivot shafts 74 and 76, respectively, and retain airfoil portions 58
in the engine
axial, circumferential and radial direction. In one form, pivot shafts 74 and
76 retain
airfoil portion 58 in outer band 50 and inner band 52 via anti-friction
bushings 82 and 84.
Anti-friction bushings 82 and 84 are operative to provide bearing surfaces for
pivot
shafts 74 and 76. Other embodiments may not include anti-friction bushings 82
and 84.
Airfoil portion 58 is operative to rotate in rotation directions 86 about
pivot axis 72.
During the operation of engine 10, air flows past vanes 54 in the general
direction
illustrated as direction 88. Vane 54 has a pressure side 90 and a suction side
92,
wherein the pressure on pressure side 90 exceeds that of suction side 92. The
pressure differential between pressure side 90 and suction side 92 may vary,
e.g.,
depending upon vane 54 camber and engine operating conditions. The pressure
differential between pressure side 90 and suction side 92 provides an impetus
to flow
from pressure side 90 to suction side 92, e.g., between airfoil portion 56 and
airfoil
portion 58. It is desirable to reduce or prevent leakage between airfoil
portion 56 and
airfoil portion 58, e.g., leakage flow from pressure side 90 to suction side
92, e.g., in
order to improve fan 12 and engine 10 efficiency. Accordingly, vanes 54
include a
sealing arrangement 94 operative to seal between airfoil portion 56 and
airfoil portion
58. Sealing arrangement 94 includes a seal strip 96 arranged to seal against
fluid flow
between airfoil portion 56 and airfoil portion 58 during the operation of
engine 10,
and to accommodate movement of one or both of airfoil portions 56 and 58,
e.g.,
rotation of airfoil portion 58 about pivot axis 72, while sealing against
fluid flow.
In one form, seal strip 96 is a rigid structure that does not substantially
deform in use or installation. In other embodiments, seal strip 96 may be a
flexible structure. In one form, seal strip 96 is formed of a polymeric
material,
such as VespelTM (commercially available from DuPont Engineering Polymers,
located in Newark, Delaware, U.S.A.) and/or TorIon TM polyamide-imide
(commercially available from Solvay Advanced Polymers, located in Alpharetta,
Georgia, U.S.A.). In other embodiments, seal strip 96 may be formed of other
materials. In one form, seal strip 96 is disposed in a groove 98. In one form,
groove 98 is disposed in a face 100 of airfoil portion 56 that faces airfoil
portion
58. In one form, seal strip 96, groove 98 and face 100 extend between tip
portion
60 and root portion 62 of airfoil portion 56. In other embodiments, seal strip
96,
groove 98 and/or face 100 may extend only partially between tip portion 60 and
root portion 62. Face 100 is formed with a radius 102 centered on pivot axis
72.
In one form, face 100 is formed integrally with airfoil portion 56. In other
embodiments, face 100 may be formed separately and affixed to airfoil portion
56. In one form, seal strip 96 is partially installed in groove 98, that is,
leaving a
portion 108 of seal strip 96 extending beyond face 100 of airfoil portion 56.
Seal
strip 96 has a width 104 greater than a width 106 of groove 98, and is
installed
into groove 98 with an interference fit, e.g., 0.001-0.002 inch. The amount of
interference may vary with the needs of the application.
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Airfoil portion 58 includes a crown 110 facing face 100 of airfoil portion 56.
In
one form, crown 110 is formed integrally with airfoil portion 58. In other
embodiments,
crown 110 may be formed separately and affixed to airfoil portion 58. Crown
110 is
formed with a radius 112 centered on pivot axis 72. In one form, crown 110
extends
between tip portion 66 and root portion 68 of airfoil portion 58, and is
positioned
opposite groove 98. In other embodiments, crown 110 may extend only partially
between tip portion 66 and root portion 68. In one form, face 100 of airfoil
portion 56 is
concave, and is operative to receive therein crown 110 opposite groove 98 in a
nested
arrangement. In other embodiments, face 100 may be convex. In one form, crown
110
of airfoil portion 58 is convex, and is operative to be received into face 100
in a nested
arrangement. In other embodiments, crown 110 may be convex, e.g., an inverted
crown. Although the depicted embodiment includes groove 98 and seal strip 96
being
located in face 100, in other embodiments, groove 98 and seal strip 96 may be
located
in crown 110.
Seal strip 96 includes a rubbing surface 114. In one form, rubbing surface 114
is
disposed opposite radius 112 of crown 110, and is operative to contact and
seal against
radius 112 of crown 110 of airfoil portion 58. During movement of airfoil
portion 58, e.g.,
when changing the camber of vane 54 by rotating airfoil portion 58 about pivot
axis 72,
rubbing surface 114 may rub against crown 110, e.g., until wear of seal strip
96
resulting from rotation of airfoil portion 58 reduces or eliminates contact
between seal
strip 96 and crown 110. In other embodiments, rubbing surface 114 may be
configured
to be in close proximity to crown 110, but without any rubbing contact. In
still other
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embodiments, seal strip 96 may be installed in crown 110, and rubbing surface
114 may
be configured to seal against face 100.
Rubbing surface 114 is preformed prior to installation into airfoil portion
56, e.g.,
machined. In one form, rubbing surface 114 is configured as a radius 116
centered
about pivot axis 72, e.g., the same radius as radius 112 of crown 110. In
other
embodiments, radius 116 may be the same radius as radius 102 of face 100 or
any
other radius suitable for the application. In still other embodiments, other
shapes for
rubbing surface 114 may be employed. In one form, rubbing surface 114 is
concave. In
other embodiments, rubbing surface 114 may take other forms, and may be, for
example, convex.
Embodiments of the present invention include a variable camber vane system for
a gas turbine engine, comprising: a first airfoil portion having a first tip
portion, a first
root portion, a face extending at least partially between the first tip
portion and the first
root portion, and a groove in the face extending at least partially between
the first tip
portion and the first root portion, wherein the groove has a groove width; a
second airfoil
portion arranged to rotate with respect to the first airfoil portion about a
pivot axis,
wherein the second airfoil portion includes a second tip portion; a second
root portion;
and a crown extending at least partially between the second tip portion and
the second
root portion, wherein the crown includes a crown radius centered about the
pivot axis
and positioned opposite the groove; and a seal strip having a seal width
greater than
the groove width and a rubbing surface opposite the crown radius, wherein the
seal strip
is at least partially disposed in the groove with an interference fit; and
wherein the seal
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strip is arranged to seal against fluid flow between the first airfoil portion
and the second
airfoil portion.
In a refinement, the seal strip is a rigid structure.
In another refinement, the rubbing surface has a rubbing surface radius the
same
as the crown radius.
In yet another refinement, the crown is formed integrally with the second
airfoil
portion.
In still another refinement, the face is formed integrally with the first
airfoil portion.
In yet still another refinement, the face is concave and operative to receive
the
crown therein.
In a further refinement, the first airfoil portion is stationary.
In a yet further refinement, the first airfoil portion and the second airfoil
portion
form at least part of an inlet guide vane having a fixed leading edge and a
variable
trailing edge; wherein the first airfoil portion includes the leading edge;
and wherein the
second airfoil portion includes the trailing edge.
In a still further refinement, the first airfoil portion and the second
airfoil portion
form at least part of an outlet guide vane having a variable leading edge and
a fixed
trailing edge; wherein the first airfoil portion includes the leading edge;
and wherein the
second airfoil portion includes the trailing edge.
Embodiments of the present invention include a gas turbine engine, comprising:
at least one of a fan and a compressor having a variable camber vane system,
the
variable camber vane system including: at least two airfoil portions adapted
to vary a
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camber of the variable camber vane system, wherein a first of the airfoil
portions
includes a groove and a second of the airfoil portions includes a crown having
a crown
radius; and a seal strip at least partially disposed in the groove with an
interference fit,
wherein the seal strip includes a rubbing surface opposite the crown radius
and
operative to seal against fluid flow between the first of the airfoil portions
and the
second of the airfoil portions.
In a refinement, the rubbing surface contacts the crown at the crown radius.
In another refinement, the seal strip is formed of a polymer material.
In yet another refinement, the seal strip is formed of at least one of Vespel
and
TorIon.
In still another refinement, the at least two airfoil portions form an inlet
guide
vane.
In a further refinement, the at least two airfoil portions form an outlet
guide vane.
Embodiments include a gas turbine engine, comprising: at least one of a fan
and
a compressor having a variable camber vane system, the variable camber vane
system
including: at least two airfoil portions adapted to vary a camber of the
variable camber
vane system, wherein a first of the airfoil portions includes a groove; and
wherein a
second of the airfoil portions includes a crown having a crown radius; and a
seal strip
disposed in the groove; wherein the seal strip has a rubbing surface radius
preformed
thereon and configured for sealing engagement with the crown.
In a refinement, the seal strip is a rigid structure formed of a polymer.
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In another refinement, the crown radius is convex, and the rubbing surface
radius
is concave.
In yet another refinement, the seal strip is fitted in the groove with an
interference
fit.
In still another refinement, the crown is nested within the first of the
airfoil
portions opposite the groove.
While the invention has been described in connection with what is presently
considered to be the most practical and preferred embodiment, it is to be
understood
that the invention is not to be limited to the disclosed embodiment(s), but on
the
contrary, is intended to cover various modifications and equivalent
arrangements
included within the spirit and scope of the appended claims, which scope is to
be
accorded the broadest interpretation so as to encompass all such modifications
and
equivalent structures as permitted under the law. Furthermore it should be
understood
that while the use of the word preferable, preferably, or preferred in the
description
above indicates that feature so described may be more desirable, it
nonetheless may
not be necessary and any embodiment lacking the same may be contemplated as
within the scope of the invention, that scope being defined by the claims that
follow. In
reading the claims it is intended that when words such as "a," "an," "at least
one" and
"at least a portion" are used, there is no intention to limit the claim to
only one item
unless specifically stated to the contrary in the claim. Further, when the
language "at
least a portion" and/or "a portion" is used the item may include a portion
and/or the
entire item unless specifically stated to the contrary.
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