U.S. patent application number 12/978843 was filed with the patent office on 2012-06-28 for gas turbine engine and variable camber vane system.
Invention is credited to Dan Molnar, James Morton, Robert A. Ress, JR..
Application Number | 20120163960 12/978843 |
Document ID | / |
Family ID | 46317012 |
Filed Date | 2012-06-28 |
United States Patent
Application |
20120163960 |
Kind Code |
A1 |
Ress, JR.; Robert A. ; et
al. |
June 28, 2012 |
GAS TURBINE ENGINE AND VARIABLE CAMBER VANE SYSTEM
Abstract
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.
Inventors: |
Ress, JR.; Robert A.;
(Carmel, IN) ; Morton; James; (Greenwood, IN)
; Molnar; Dan; (Lebanon, IN) |
Family ID: |
46317012 |
Appl. No.: |
12/978843 |
Filed: |
December 27, 2010 |
Current U.S.
Class: |
415/173.1 ;
415/208.1 |
Current CPC
Class: |
F04D 29/544 20130101;
F01D 17/162 20130101; F04D 29/563 20130101 |
Class at
Publication: |
415/173.1 ;
415/208.1 |
International
Class: |
F01D 9/02 20060101
F01D009/02; F01D 25/24 20060101 F01D025/24; F01D 5/20 20060101
F01D005/20 |
Goverment Interests
GOVERNMENT RIGHTS
[0001] 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.
Claims
1. 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 strip is arranged to
seal against fluid flow between the first airfoil portion and the
second airfoil portion.
2. The variable camber vane system of claim 1, wherein the seal
strip is a rigid structure.
3. The variable camber vane system of claim 1, wherein the rubbing
surface has a rubbing surface radius the same as the crown
radius.
4. The variable camber vane system of claim 1, wherein the crown is
formed integrally with the second airfoil portion.
5. The variable camber vane system of claim 1, wherein the face is
formed integrally with the first airfoil portion.
6. The variable camber vane system of claim 1, wherein the face is
concave and operative to receive the crown therein.
7. The variable camber vane system of claim 1, wherein the first
airfoil portion is stationary.
8. The variable camber vane system of claim 7, wherein 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.
9. The variable camber vane system of claim 7, wherein 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.
10. 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 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.
11. The gas turbine engine of claim 10, wherein the rubbing surface
contacts the crown at the crown radius.
12. The gas turbine engine of claim 10, wherein the seal strip is
formed of a polymer material.
13. The gas turbine engine of claim 12, wherein the seal strip is
formed of at least one of Vespel and Torlon.
14. The gas turbine engine of claim 10, wherein the at least two
airfoil portions form an inlet guide vane.
15. The gas turbine engine of claim 10, wherein the at least two
airfoil portions form an outlet guide vane.
16. 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.
17. The gas turbine engine of claim 16, wherein the seal strip is a
rigid structure formed of a polymer.
18. The gas turbine engine of claim 16, wherein the crown radius is
convex, and wherein the rubbing surface radius is concave.
19. The gas turbine engine of claim 16, wherein the seal strip is
fitted in the groove with an interference fit.
20. The gas turbine engine of claim 16, wherein the crown is nested
within the first of the airfoil portions opposite the groove.
Description
FIELD OF THE INVENTION
[0002] The present invention relates to gas turbine engines, and
more particularly, to gas turbine engines with variable camber vane
systems.
BACKGROUND
[0003] 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.
SUMMARY
[0004] 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.
BRIEF DESCRIPTION OF THE DRAWINGS
[0005] The description herein makes reference to the accompanying
drawings wherein like reference numerals refer to like parts
throughout the several views, and wherein:
[0006] 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.
[0007] 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.
[0008] 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.
[0009] FIG. 4 depicts some aspects of the variable camber guide
vane system of FIG. 3.
[0010] FIG. 5 depicts some aspects of a non-limiting example of a
seal strip in accordance with an embodiment of the present
invention.
DETAILED DESCRIPTION
[0011] 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.
[0012] 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.
[0013] 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 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.
[0014] 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 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.
[0015] 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.
[0016] 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 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.
[0017] 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 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.
[0018] 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.
[0019] 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.
[0020] 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
Vespel (commercially available from DuPont Engineering Polymers,
located in Newark, Del., U.S.A.) and/or Torlon polyamide-imide
(commercially available from Solvay Advanced Polymers, located in
Alpharetta, Ga., 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.
[0021] 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.
[0022] 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 embodiments,
seal strip 96 may be installed in crown 110, and rubbing surface
114 may be configured to seal against face 100.
[0023] 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.
[0024] 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 strip is arranged to seal against fluid flow
between the first airfoil portion and the second airfoil
portion.
[0025] In a refinement, the seal strip is a rigid structure.
[0026] In another refinement, the rubbing surface has a rubbing
surface radius the same as the crown radius.
[0027] In yet another refinement, the crown is formed integrally
with the second airfoil portion.
[0028] In still another refinement, the face is formed integrally
with the first airfoil portion.
[0029] In yet still another refinement, the face is concave and
operative to receive the crown therein.
[0030] In a further refinement, the first airfoil portion is
stationary.
[0031] 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.
[0032] 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.
[0033] 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 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.
[0034] In a refinement, the rubbing surface contacts the crown at
the crown radius.
[0035] In another refinement, the seal strip is formed of a polymer
material.
[0036] In yet another refinement, the seal strip is formed of at
least one of Vespel and Torlon.
[0037] In still another refinement, the at least two airfoil
portions form an inlet guide vane.
[0038] In a further refinement, the at least two airfoil portions
form an outlet guide vane.
[0039] 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.
[0040] In a refinement, the seal strip is a rigid structure formed
of a polymer.
[0041] In another refinement, the crown radius is convex, and the
rubbing surface radius is concave.
[0042] In yet another refinement, the seal strip is fitted in the
groove with an interference fit.
[0043] In still another refinement, the crown is nested within the
first of the airfoil portions opposite the groove.
[0044] 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.
* * * * *