U.S. patent application number 11/752945 was filed with the patent office on 2009-03-12 for variable area turbine vane arrangement.
Invention is credited to George T. Suljak, JR..
Application Number | 20090067978 11/752945 |
Document ID | / |
Family ID | 40432022 |
Filed Date | 2009-03-12 |
United States Patent
Application |
20090067978 |
Kind Code |
A1 |
Suljak, JR.; George T. |
March 12, 2009 |
VARIABLE AREA TURBINE VANE ARRANGEMENT
Abstract
A turbine section of a gas turbine engine includes a ring of
turbine nozzle segments each having paired turbine vanes. Turbine
throat area is modulated by rotating each rotational turbine vanes
about an axis of rotation which is located such that rotation
changes the turbine throat area concurrently between one rotational
stator vane and two adjacent fixed turbine vanes. Each paired
turbine vane doublet includes at least one rotational turbine vane
between two fixed turbine vanes.
Inventors: |
Suljak, JR.; George T.;
(Vemon, CT) |
Correspondence
Address: |
CARLSON, GASKEY & OLDS/PRATT & WHITNEY
400 WEST MAPLE ROAD, SUITE 350
BIRMINGHAM
MI
48009
US
|
Family ID: |
40432022 |
Appl. No.: |
11/752945 |
Filed: |
May 24, 2007 |
Current U.S.
Class: |
415/1 ;
415/160 |
Current CPC
Class: |
F01D 17/162
20130101 |
Class at
Publication: |
415/1 ;
415/160 |
International
Class: |
F04D 29/56 20060101
F04D029/56 |
Claims
1. A turbine nozzle segment for a gas turbine engine comprising: a
fixed turbine vane; and a rotational turbine vane adjacent said
fixed turbine vane, said rotational turbine vane rotatable about an
axis of rotation relative said fixed turbine vane.
2. The turbine nozzle segment as recited in claim 1, wherein said
fixed turbine vane and said rotational turbine vane are located
between an outer vane platform segment and an inner vane platform
segment.
3. The turbine nozzle segment as recited in claim 2, wherein said
fixed turbine vane is fixed to said outer vane platform segment and
said inner vane platform segment.
4. The turbine nozzle segment as recited in claim 1, further
comprising an actuator system having an actuator rod mounted
through said rotational turbine vane at said axis of rotation.
5. The turbine nozzle segment as recited in claim 4, wherein said
rotational turbine vane is rotated by a unison ring driven
linkage.
6. The turbine nozzle segment as recited in claim 1, wherein said
axis of rotation is aft of a geometric center of gravity of a cross
section of said rotational turbine vane.
7. The turbine nozzle segment as recited in claim 1, wherein said
axis of rotation is located approximately midway between a trailing
edge of said fixed turbine vane and a trailing edge of said
rotational turbine vane.
8. A turbine section of a gas turbine engine comprising: an annular
outer vane platform; an annular inner vane platform; a fixed
turbine vane fixed to said annular outer vane platform and said
annular inner vane platform; and a rotational turbine vane between
said annular outer vane platform and said annular inner vane
platform, said rotational turbine vane rotatable about an axis of
rotation aft of a geometric center of gravity of a cross section of
said rotational turbine vane.
9. The turbine section as recited in claim 8, wherein said annular
outer vane platform and said annular inner vane platform are formed
from a respective multiple of outer vane platform segments and a
multiple of inner vane platform segments, wherein each of said
multiple of outer vane platform segments and said multiple of inner
vane platform segments include at least one of said fixed turbine
vanes.
10. The turbine section as recited in claim 9, wherein each of said
multiple of outer vane platform segments and said multiple of inner
vane platform segments include at least one of said rotational
turbine vane between a first fixed turbine vane and a second fixed
turbine vane.
11. The turbine section as recited in claim 8, wherein said fixed
turbine vane alternates with said rotational turbine vane.
12. The turbine section as recited in claim 8, wherein said axis of
rotation is located approximately midway between a trailing edge of
said fixed turbine vane and a trailing edge of said rotational
turbine vane
13. A method of varying a turbine nozzle throat area of a gas
turbine engine comprising the steps of: (A) locating a rotational
turbine vane between a first fixed turbine vane and a second fixed
turbine vane; and (B) rotating the rotational turbine vane about an
axis of rotation to vary a throat area concurrently between the
rotational stator vane and both the first fixed turbine vane and
the second fixed turbine vane.
14. A method as recited in claim 13, wherein said step (A) further
comprises: (a) alternating each of a multiple of rotational turbine
vanes with a multiple of fixed turbine vanes about a turbine
section of the gas turbine engine.
15. A method as recited in claim 14, wherein said step (B) further
comprises: (a) rotating a unison ring to simultaneously rotate each
of the multiple of rotational turbine vanes.
16. A method as recited in claim 13, wherein said step (B) further
comprises: (a) locating the axis of rotation aft of a geometric
center of gravity of a cross section of the rotational turbine
vane.
Description
BACKGROUND OF THE INVENTION
[0001] The present invention relates to a gas turbine engine
turbine section, and more particularly to a ring of paired turbine
vanes doublets in which one vane of each doublet rotates to
modulate turbine throat area.
[0002] The core engine of a gas turbine engine typically includes a
multistage axial compressor which provides compressed air to a
combustor wherein it is mixed with fuel and ignited for generating
hot combustion gas which flows downstream through a high pressure
turbine nozzle and in turn through one or more stages of turbine
rotor blades. The high pressure turbine blades are joined to a
rotor disk which is joined to the compressor by a corresponding
drive shaft, with the turbine blades extracting energy for powering
the compressor during operation. In a two spool engine, a second
shaft joins a fan upstream of the compressor to a low pressure
turbine disposed downstream from the high pressure turbine.
[0003] Typical turbine nozzles, such as high pressure and low
pressure turbine nozzles, have fixed vane configurations and fixed
turbine nozzle throat areas. Variable cycle engines are being
developed to maximize performance and efficiency over subsonic and
supersonic flight conditions. Some engines provide variability in
compressor turbine vanes by mounting each vane on a radial spindle
and collectively rotating each row of compressor vanes using an
annular unison ring attached to corresponding lever arms joined to
each of the spindles. Each compressor vane rotates about a radial
axis, with suitable hub and tip clearances which permit rotation of
the vanes.
[0004] Although it would be desirable to obtain variable flow
through turbine nozzles by adjusting the throat areas thereof,
previous attempts thereat have proved difficult because of severe
operating environment of the turbine nozzles. The severe
temperature environment of the turbine nozzle typically requires
suitable cooling of the individual vanes, with differential
temperature gradients. Nozzle vanes are also subject to substantial
aerodynamic loads from the combustion gas during operation.
Furthermore, adjustable turbine nozzle vanes may reduce the
structural integrity and durability of the nozzle segments in view
of the increased degree of freedom therebetween.
[0005] Accordingly, it is desirable to provide a variable area
turbine nozzle having a relatively uncomplicated rotation, support
and sealing structure to provide variable nozzle throat area
capability yet minimize turbine pressure loss, leakage, expense and
weight.
SUMMARY OF THE INVENTION
[0006] The turbine section of a gas turbine engine according to the
present invention provides a ring of turbine nozzle segments each
having paired turbine vanes (Doublets). Turbine throat area is
modulated by rotating one of the doublet vanes about an axis
located towards the rotatable turbine vane trailing edge. The vane
adjacent each rotational turbine vane is a fixed turbine vane which
provides a rigid structure to support the rotational turbine vanes
which thereby form only a portion of the turbine section. The
center of rotation of the rotational turbine vane is located such
that rotation of the rotational turbine vane changes the turbine
throat area concurrently between the rotational stator vane and
both adjacent fixed turbine vanes.
[0007] The present invention therefore provides a variable area
turbine nozzle having relatively uncomplicated rotation, support
and sealing structure to provide variable nozzle throat area
capability yet minimize turbine pressure loss, leakage, expense and
weight.
BRIEF DESCRIPTION OF THE DRAWINGS
[0008] The various features and advantages of this invention will
become apparent to those skilled in the art from the following
detailed description of the currently disclosed embodiment. The
drawings that accompany the detailed description can be briefly
described as follows:
[0009] FIG. 1 is a general perspective view an exemplary gas
turbine engine embodiment for use with the present invention;
[0010] FIG. 2 is an expanded view of a vane portion of one turbine
stage within a turbine section of the gas turbine engine, the vane
portion formed from a multiple of turbine nozzle segments;
[0011] FIG. 3 is an expanded partial phantom view of one variable
turbine nozzle segment; and
[0012] FIG. 4 is a top schematic representation of the throat
change performed by the turbine section according to the present
invention.
DETAILED DESCRIPTION OF THE DISCLOSED EMBODIMENT
[0013] FIG. 1 schematically illustrates a gas turbine engine 10
which generally includes a fan section 12, a compressor section 14,
a combustor section 16, a turbine section 18, and a nozzle section
20 along a longitudinal axis X. The gas turbine engine 10 of the
disclosed embodiment is a relatively low bypass gas turbine engine.
Although the disclosed embodiment illustrates a 3-stage fan, a
6-stage compressor, an annular combustor, a single stage
high-pressure turbine, and a 2-stage low pressure turbine, various
other gas turbine engines will benefit from the present
invention.
[0014] The engine 10 is configured to provide a variable area
turbine nozzle to selectively control the flow of the combustion
gas 12 from the combustor section 16 through the turbine section
18. The engine 10 is also referred to as including a Controlled
Area Turbine Nozzle (CATN).
[0015] Referring to FIG. 2, a turbine nozzle segment 30 includes an
arcuate outer vane platform segment 32 and an arcuate inner vane
platform segment 34 radially spaced apart from each other. The
arcuate outer vane platform segment 32 may form a portion of an
outer core engine structure 22 and the arcuate inner vane platform
segment 34 may form a portion of an inner core engine structure 24
(FIG. 1) to at least partially define an annular turbine nozzle
core gas flow path 26.
[0016] The circumferentially adjacent vane platform segments 32, 34
define split lines 36 which thermally uncouple adjacent turbine
nozzle segments 30 which may be conventionally sealed therebetween,
with, for example only, spline seals. That is, the temperature
environment of the turbine section 18 and the substantial
aerodynamic and thermal loads are accommodated by the plurality of
circumferentially adjoining nozzle segments 30 which collectively
form a full, annular ring about the centerline axis X of the
engine.
[0017] Each turbine nozzle segment 30 includes a multiple (two
shown) of circumferentially spaced apart turbine vanes 38, 40 which
extend radially between the vane platform segments 32, 34. In the
disclosed embodiment, each nozzle segment 30 includes one fixed
turbine vane 38 and one rotational turbine vane 40 (doublet)
between the vane platform segments 32, 34 to provide a rigid
structural assembly which accommodates thermal and aerodynamic
loads during operation. That is, the full, annular ring formed by
the multiple of turbine nozzle segments 30 provide a vane portion
of one stage in the turbine section 18 which is defined by the
alternating fixed and rotational turbine vanes 38, 40.
[0018] Each turbine nozzle segment 30 includes at least one fixed
turbine vane 38 and at least one rotational turbine vane 40 such
that the fixed turbine vane 38 and the vane platform segments 32,
34 form a box structure. The vane platform segments 32, 34 may
include features 50 to mount each nozzle segment 30 to other engine
static structures. It should be understood that although the
illustrated embodiment discloses a doublet arrangement, any number
of fixed turbine vanes 38 and rotational turbine vanes 40 may be
provided in each turbine nozzle segment 30. Movement of the
rotational turbine vanes 40 relative the adjacent fixed turbine
vanes 38 effectuates a change in throat area formed by the ring of
nozzle segments 30 as will be further described below.
[0019] Referring to FIG. 3, each turbine vane 38, 40 includes a
respective airfoil portion 42F, 42R defined by an outer airfoil
wall surface 44F 44R between the leading edge 46F, 46R and a
trailing edge 48F, 48R. Each turbine vane 38, 40 may include a
fillet 52 to provide a transition between the airfoil portion 42F,
42R and the vane platform segments 32, 34. The outer airfoil wall
surface 44 is typically shaped for use in, for example only, a
first stage, or other stage, of a high pressure and low pressure
stage of the turbine section. The outer airfoil wall 44F, 44R
typically have a generally concave shaped portion forming a
pressure side 44FP, 44RP and a generally convex shaped portion
forming a suction side 44FS, 44RS. It should be understood that
respective airfoil portion 42F, 42R defined by the outer airfoil
wall surface 44F 44R may be generally equivalent or separately
tailored to optimize flow characteristics and transient thermal
expansion issues.
[0020] An actuator system 54 includes an actuator such as an outer
diameter unison ring (illustrated schematically at 56) which
rotates an actuator arm 58 and an actuator rod 60 which passes
through the inner vane platform segment 32, the rotational turbine
vane 40, and the outer vane platform segment 34. The actuator rod
60 rotates each rotational turbine vane 40 about a vane axis of
rotation 62 relative the adjacent fixed turbine vanes 38 to
selectively vary the turbine nozzle throat area. Since the fixed
turbine vane 38 and vane platform segments 32, 34 provide a rigid
structure, the rotational turbine vane 40 may include a relatively
less complicated rotation, support and sealing structure to provide
the variable nozzle throat area capability which minimizes turbine
pressure loss, leakage, expense and weight.
[0021] The vane axis of rotation 62 is located approximately midway
between the trailing edges of an adjacent fixed turbine vanes 38
and rotational turbine vane 40 to close the throat area between the
rotational turbine vane 40 and the adjacent fixed turbine vanes 38
on either side of the rotational turbine vane 40 simultaneously
(FIG. 4). Airfoils are conventionally rotated around the geometric
center of gravity (CG) of the airfoil cross section. Here, the
rotational turbine vane 40 vane axis of rotation 62 is biased
toward the trailing edge 48R of the rotational turbine vane 40. In
one embodiment, a distance L between the trailing edges of an
adjacent fixed turbine vanes 38 and rotational turbine vane 40, may
be about 1.6 inches (FIG. 4). The rotational turbine vane 40 axis
of rotation 62 is then positioned at L/2 or about 0.8 inches from
each adjacent fixed turbine vane 38 such that the axis of rotation
62 is located axially aft of the conventional geometric CG.
[0022] In operation, rotation of the rotational turbine vanes 40
between a nominal position and a rotated position selectively
changes the turbine nozzle throat area as each rotational turbine
vane 40 concurrently changes the throat area between itself and
both adjacent fixed turbine vanes 38. Since only half the vanes are
rotated, the required rotation is less since rotation changes the
throat on both sides simultaneously, with less change in the gas
exit angle directed to the turbine blades. Furthermore, since only
half of the vanes are rotated, the complexity and load requirements
of the actuator system 54 are reduced. The alternating
rotational-fixed vane arrangement also facilitates a relatively
less complicated rotation, support and sealing structure to provide
the variable nozzle throat area capability to minimize turbine
pressure loss, leakage, expense and weight.
[0023] It should be understood that relative positional terms such
as "forward," "aft," "upper," "lower," "above," "below," and the
like are with reference to the normal operational attitude of the
device and should not be considered otherwise limiting.
[0024] It should be understood that although a particular component
arrangement is disclosed in the illustrated embodiment, other
arrangements will benefit from the instant invention.
[0025] Although particular step sequences are shown, described, and
claimed, it should be understood that steps may be performed in any
order, separated or combined unless otherwise indicated and will
still benefit from the present invention.
[0026] The foregoing description is exemplary rather than defined
by the limitations within. Many modifications and variations of the
present invention are possible in light of the above teachings. The
disclosed embodiments of this invention have been disclosed,
however, one of ordinary skill in the art would recognize that
certain modifications would come within the scope of this
invention. It is, therefore, to be understood that within the scope
of the appended claims, the invention may be practiced otherwise
than as specifically described. For that reason the following
claims should be studied to determine the true scope and content of
this invention.
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