U.S. patent application number 14/783965 was filed with the patent office on 2016-03-10 for stator vane platform with flanges.
The applicant listed for this patent is UNITED TECHNOLOGIES CORPORATION. Invention is credited to Shelton O. Duelm, Kevin L. Rugg.
Application Number | 20160069199 14/783965 |
Document ID | / |
Family ID | 52468770 |
Filed Date | 2016-03-10 |
United States Patent
Application |
20160069199 |
Kind Code |
A1 |
Duelm; Shelton O. ; et
al. |
March 10, 2016 |
STATOR VANE PLATFORM WITH FLANGES
Abstract
Disclosed is a gas turbine engine that includes a plurality of
stator vanes. Each of the stator vanes includes an airfoil section
and a platform provided at an end of the airfoil section. The
platforms are provided with flanges on opposed ends thereof, and
the flanges are fastened to corresponding flanges of an adjacent
one of the plurality of stator vanes.
Inventors: |
Duelm; Shelton O.;
(Wethersfield, CT) ; Rugg; Kevin L.; (Fairfield,
CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
UNITED TECHNOLOGIES CORPORATION |
Hartford |
CT |
US |
|
|
Family ID: |
52468770 |
Appl. No.: |
14/783965 |
Filed: |
April 11, 2014 |
PCT Filed: |
April 11, 2014 |
PCT NO: |
PCT/US2014/033719 |
371 Date: |
October 12, 2015 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
|
|
61811149 |
Apr 12, 2013 |
|
|
|
Current U.S.
Class: |
415/209.3 |
Current CPC
Class: |
Y02T 50/672 20130101;
F01D 9/042 20130101; Y02T 50/673 20130101; F01D 25/246 20130101;
Y02T 50/60 20130101 |
International
Class: |
F01D 9/04 20060101
F01D009/04; F01D 25/24 20060101 F01D025/24 |
Claims
1. A gas turbine engine, comprising: a plurality of stator vanes,
each of the plurality of stator vanes including an airfoil section
and a platform provided at an end of the airfoil section, the
platform provided with flanges on opposed ends thereof, the flanges
being fastened to corresponding flanges of an adjacent one of the
plurality of stator vanes.
2. The gas turbine engine as recited in claim 1, wherein the
flanges are provided on circumferential ends of the platform.
3. The gas turbine engine as recited in claim 2, wherein a length
dimension of the flanges is provided in a direction generally
parallel to an engine central longitudinal axis.
4. The gas turbine engine as recited in claim 3, wherein the
flanges include inner and outer faces generally perpendicular the
platform.
5. The gas turbine engine as recited in claim 4, wherein the
flanges are fastened to one another such that the outer faces of
the flanges directly abut one another.
6. The gas turbine engine as recited in claim 5, wherein the
flanges include openings configured to receive fasteners.
7. The gas turbine engine as recited in claim 6, wherein the
fasteners include at least one of bolts, pins, and springs.
8. The gas turbine engine as recited in claim 6, wherein the
fasteners fasten adjacent flanges to one another.
9. The gas turbine engine as recited in claim 6, wherein the
fasteners fasten the flanges to a static structure of the gas
turbine engine via a respective coupling.
10. The gas turbine engine as recited in claim 6, including at
least one resilient member connected to the flange and to the
platform to reduce bending of the flange.
11. The gas turbine engine as recited in claim 1, wherein at least
one of the plurality of stator vanes is made of a ceramic matrix
composite (CMC) material.
12. The gas turbine engine as recited in claim 1, further including
a compressor section, a combustor section, and a turbine section,
the turbine section including a stationary stage, the stator vanes
provided in the stationary stage.
13. The gas turbine engine as recited in claim 1, wherein each of
the plurality of stator vanes includes a first platform at a
radially outer end of the airfoil section, and a second platform at
a radially inner end of the airfoil section, each of the first and
second platforms provided with flanges on opposed ends thereof,
wherein the flanges associated with the first platform are fastened
to flanges of a first platform of an adjacent stator vane, and the
flanges associated with the second platform are not fastened to the
flanges of an adjacent stator vane to allow the second platform to
be radially free.
14. A stator vane for a gas turbine engine, comprising: an airfoil
section; a platform provided at an end of the airfoil section, the
platform having a fore end, an aft end, and circumferential ends,
the platform including flanges provided on the circumferential
ends, the flanges configured to abut flanges of a similar stator
vane.
15. The stator vane as recited in claim 14, wherein the airfoil
section is provided generally along a radial direction.
16. The stator vane as recited in claim 15, wherein the flanges are
provided such that a length thereof is generally perpendicular to
the radial direction.
17. The stator vane as recited in claim 14, wherein each flange
includes at least one opening for receiving a fastener.
18. The stator vane as recited in claim 14, wherein the stator vane
is a singlet.
19. The stator vane as recited in claim 14, wherein at least a
portion of the stator vane is made of ceramic matrix composite
(CMC) material.
20. The stator vane as recited in claim 14, wherein the flanges are
integrally formed with the stator vane.
Description
BACKGROUND
[0001] Gas turbine engines typically include a compressor section,
a combustor section and a turbine section. During operation, air is
pressurized in the compressor section and is mixed with fuel and
burned in the combustor section to generate hot combustion gases.
The hot combustion gases are communicated through the turbine
section, which extracts energy from the hot combustion gases to
power the compressor section and other gas turbine engine
loads.
[0002] Both the compressor and turbine sections may include
alternating series of rotating blades and stationary vanes that
extend into the core flow path of the gas turbine engine. For
example, in the turbine section, turbine blades rotate and extract
energy from the hot combustion gases that are communicated along
the core flow path of the gas turbine engine. The turbine vanes,
which generally do not rotate, guide the airflow and prepare it for
the next set of blades.
[0003] In turbine vane design, there is an emphasis on
stress-resistant airfoil and platform designs, with reduced losses,
increased lift and turning efficiency, and improved turbine
performance and service life.
SUMMARY
[0004] In one exemplary embodiment, a gas turbine engine includes a
plurality of stator vanes. Each of the plurality of stator vanes
includes an airfoil section and a platform provided at an end of
the airfoil section. The platform is provided with flanges on
opposed ends thereof, and the flanges re fastened to corresponding
flanges of an adjacent one of the plurality of stator vanes.
[0005] In a further embodiment of any of the above, the flanges are
provided on circumferential ends of the platform.
[0006] In a further embodiment of any of the above, a length
dimension of the flanges is provided in a direction generally
parallel to an engine central longitudinal axis.
[0007] In a further embodiment of any of the above, the flanges
include inner and outer faces generally perpendicular the
platform.
[0008] In a further embodiment of any of the above, the flanges are
fastened to one another such that the outer faces of the flanges
directly abut one another.
[0009] In a further embodiment of any of the above, the flanges
include openings configured to receive fasteners.
[0010] In a further embodiment of any of the above, the fasteners
include at least one of bolts, pins, and springs.
[0011] In a further embodiment of any of the above, the fasteners
fasten adjacent flanges to one another.
[0012] In a further embodiment of any of the above, the fasteners
fasten the flanges to a static structure of the gas turbine engine
via a respective coupling.
[0013] In a further embodiment of any of the above, the gas turbine
engine includes at least one resilient member connected to the
flange and to the platform to reduce bending of the flange.
[0014] In a further embodiment of any of the above, at least one of
the plurality of stator vanes is made of a ceramic matrix composite
(CMC) material.
[0015] In a further embodiment of any of the above, each of the
plurality of stator vanes includes a first platform at a radially
outer end of the airfoil section, and a second platform at a
radially inner end of the airfoil section, each of the first and
second platforms provided with flanges on opposed ends thereof,
wherein the flanges associated with the first platform are fastened
to flanges of a first platform of an adjacent stator vane, and the
flanges associated with the second platform are not fastened to the
flanges of an adjacent stator vane to allow the second platform to
be radially free.
[0016] In a further embodiment of any of the above, the gas turbine
engine includes a compressor section, a combustor section, and a
turbine section, the turbine section including a stationary stage,
the stator vanes provided in the stationary stage.
[0017] In another exemplary embodiment, a stator vane for a gas
turbine engine includes an airfoil section and a platform provided
at an end of the airfoil section. The platform has a fore end, an
aft end, and circumferential ends. The platform further includes
flanges provided on the circumferential ends, and the flanges are
configured to abut flanges of a similar stator vane.
[0018] In a further embodiment of any of the above, the airfoil
section is provided generally along a radial direction.
[0019] In a further embodiment of any of the above, the flanges are
provided such that a length thereof is generally perpendicular to
the radial direction.
[0020] In a further embodiment of any of the above, each flange
includes at least one opening for receiving a fastener.
[0021] In a further embodiment of any of the above, the stator vane
is a singlet.
[0022] In a further embodiment of any of the above, at least a
portion of the stator vane is made of ceramic matrix composite
(CMC) material.
[0023] In a further embodiment of any of the above, the flanges are
integrally formed with the stator vane.
[0024] These and other features of the present disclosure can be
best understood from the following drawings and detailed
description.
BRIEF DESCRIPTION OF THE DRAWINGS
[0025] The drawings can be briefly described as follows:
[0026] FIG. 1 schematically illustrates a gas turbine engine
embodiment.
[0027] FIG. 2 is a cross-sectional view through a high pressure
turbine section.
[0028] FIG. 3 is a perspective view of an example stator vane.
[0029] FIG. 4 is an example stage of the gas turbine engine,
showing a plurality of stator vanes.
[0030] FIGS. 5A-5B schematically illustrate examples for fastening
adjacent stator vanes to one another.
DETAILED DESCRIPTION
[0031] FIG. 1 schematically illustrates an example gas turbine
engine 20 that includes a fan section 22, a compressor section 24,
a combustor section 26 and a turbine section 28. Alternative
engines might include an augmenter section (not shown) among other
systems or features. The fan section 22 drives air along a bypass
flow path B while the compressor section 24 draws air in along a
core flow path C where air is compressed and communicated to a
combustor section 26. In the combustor section 26, air is mixed
with fuel and ignited to generate a high pressure exhaust gas
stream that expands through the turbine section 28 where energy is
extracted and utilized to drive the fan section 22 and the
compressor section 24.
[0032] Although the disclosed non-limiting embodiment depicts a
turbofan gas turbine engine, it should be understood that the
concepts described herein are not limited to use with turbofans as
the teachings may be applied to other types of turbine engines; for
example a turbine engine including a three-spool architecture in
which three spools concentrically rotate about a common axis and
where a low spool enables a low pressure turbine to drive a fan via
a gearbox, an intermediate spool that enables an intermediate
pressure turbine to drive a first compressor of the compressor
section, and a high spool that enables a high pressure turbine to
drive a high pressure compressor of the compressor section. The
concepts disclosed herein can further be applied outside of gas
turbine engines, such as in the context of wind turbines.
[0033] The example engine 20 generally includes a low speed spool
30 and a high speed spool 32 mounted for rotation about an engine
central longitudinal axis X relative to an engine static structure
36 via several bearing systems 38. It should be understood that
various bearing systems 38 at various locations may alternatively
or additionally be provided.
[0034] The low speed spool 30 generally includes an inner shaft 40
that connects a fan 42 and a low pressure (or first) compressor
section 44 to a low pressure (or first) turbine section 46. The
inner shaft 40 drives the fan 42 through a speed change device,
such as a geared architecture 48, to drive the fan 42 at a lower
speed than the low speed spool 30. The high-speed spool 32 includes
an outer shaft 50 that interconnects a high pressure (or second)
compressor section 52 and a high pressure (or second) turbine
section 54. The inner shaft 40 and the outer shaft 50 are
concentric and rotate via the bearing systems 38 about the engine
central longitudinal axis X.
[0035] A combustor 56 is arranged between the high pressure
compressor 52 and the high pressure turbine 54. In one example, the
high pressure turbine 54 includes at least two stages to provide a
double stage high pressure turbine 54. In another example, the high
pressure turbine 54 includes only a single stage. As used herein, a
"high pressure" compressor or turbine experiences a higher pressure
than a corresponding "low pressure" compressor or turbine.
[0036] The example low pressure turbine 46 has a pressure ratio
that is greater than about five (5). The pressure ratio of the
example low pressure turbine 46 is measured prior to an inlet of
the low pressure turbine 46 as related to the pressure measured at
the outlet of the low pressure turbine 46 prior to an exhaust
nozzle.
[0037] A mid-turbine frame 57 of the engine static structure 36 is
arranged generally between the high pressure turbine 54 and the low
pressure turbine 46. The mid-turbine frame 57 further supports
bearing systems 38 in the turbine section 28 as well as setting
airflow entering the low pressure turbine 46.
[0038] The core airflow C is compressed by the low pressure
compressor 44 then by the high pressure compressor 52 mixed with
fuel and ignited in the combustor 56 to produce high speed exhaust
gases that are then expanded through the high pressure turbine 54
and low pressure turbine 46. The mid-turbine frame 57 includes
vanes 59, which are in the core airflow path and function as an
inlet guide vane for the low pressure turbine 46. Utilizing the
vane 59 of the mid-turbine frame 57 as the inlet guide vane for low
pressure turbine 46 decreases the length of the low pressure
turbine 46 without increasing the axial length of the mid-turbine
frame 57. Reducing or eliminating the number of vanes in the low
pressure turbine 46 shortens the axial length of the turbine
section 28. Thus, the compactness of the gas turbine engine 20 is
increased and a higher power density may be achieved.
[0039] The disclosed gas turbine engine 20 in one example is a
high-bypass geared aircraft engine. In a further example, the gas
turbine engine 20 includes a bypass ratio greater than about six
(6), with an example embodiment being greater than about ten (10).
The example geared architecture 48 is an epicyclical gear train,
such as a planetary gear system, star gear system or other known
gear system, with a gear reduction ratio of greater than about
2.3.
[0040] In one disclosed embodiment, the gas turbine engine 20
includes a bypass ratio greater than about ten (10:1) and the fan
diameter is significantly larger than an outer diameter of the low
pressure compressor 44. It should be understood, however, that the
above parameters are only exemplary of one embodiment of a gas
turbine engine including a geared architecture and that the present
disclosure is applicable to other gas turbine engines.
[0041] A significant amount of thrust is provided by the bypass
flow B due to the high bypass ratio. The fan section 22 of the
engine 20 is designed for a particular flight condition--typically
cruise at about 0.8 Mach and about 35,000 feet. The flight
condition of 0.8 Mach and 35,000 ft., with the engine at its best
fuel consumption--also known as "bucket cruise Thrust Specific Fuel
Consumption (`TSFC`)"--is the industry standard parameter of
pound-mass (lbm) of fuel per hour being burned divided by
pound-force (lbf) of thrust the engine produces at that minimum
point.
[0042] "Low fan pressure ratio" is the pressure ratio across the
fan blade alone, without a Fan Exit Guide Vane ("FEGV") system. The
low fan pressure ratio as disclosed herein according to one
non-limiting embodiment is less than about 1.50. In another
non-limiting embodiment the low fan pressure ratio is less than
about 1.45.
[0043] "Low corrected fan tip speed" is the actual fan tip speed in
ft/sec divided by an industry standard temperature correction of
[(Tram.degree. R)/(518.7.degree. R)].sup.0.5. The "Low corrected
fan tip speed", as disclosed herein according to one non-limiting
embodiment, is less than about 1150 ft/second.
[0044] The example gas turbine engine includes the fan 42 that
comprises in one non-limiting embodiment less than about twenty-six
(26) fan blades. In another non-limiting embodiment, the fan
section 22 includes less than about twenty (20) fan blades.
Moreover, in one disclosed embodiment the low pressure turbine 46
includes no more than about six (6) turbine rotors schematically
indicated at 34. In another non-limiting example embodiment the low
pressure turbine 46 includes about three (3) turbine rotors. A
ratio between the number of fan blades 42 and the number of low
pressure turbine rotors is between about 3.3 and about 8.6. The
example low pressure turbine 46 provides the driving power to
rotate the fan section 22 and therefore the relationship between
the number of turbine rotors 34 in the low pressure turbine 46 and
the number of blades 42 in the fan section 22 disclose an example
gas turbine engine 20 with increased power transfer efficiency.
[0045] Referring to FIG. 2, a cross-sectional view through a high
pressure turbine section 54 is illustrated. In the example high
pressure turbine section 54, a first stage array 54a of
circumferentially spaced fixed vanes 60 is included. A first stage
array 54b of circumferentially spaced turbine blades 64, mounted to
a rotor disk 68, is arranged axially downstream of the first fixed
vane array 54a.
[0046] The turbine blades each include a tip 80 adjacent to a blade
outer air seal 70 of a case structure 72. The first stage array 54b
of turbine blades are arranged within a core flow path C and are
operatively connected to a spool 32.
[0047] Each vane 60 includes an inner platform 74 and an outer
platform 76 respectively defining inner and outer flow paths. The
platforms 74, 76 are interconnected by an airfoil section 78
extending in a radial direction Z. It should be understood that the
turbine vanes may be discrete from one another or arranged in
integrated clusters. The airfoil section 78 includes leading and
trailing edges 82, 84.
[0048] The airfoil section 78 is provided between pressure
(concave) and suction (convex) sides 85, 87 in an airfoil thickness
direction, which is generally perpendicular to a chord-wise
direction provided between the leading and trailing edges 82, 84
(FIG. 3). Multiple turbine vanes 60 are arranged circumferentially
in a circumferential direction R. The airfoil section 78 typically
includes multiple cooling holes (not shown).
[0049] In one example, the first stage array 54a consists of 32
turbine vanes 60, but the number may vary according to engine size.
The turbine vanes 60 in one example are formed of a ceramic matrix
composite (CMC) material. As is known in this art, a CMC material
is one that includes fibers (such as carbon, silicon or glass
fibers, as example) supported within a ceramic matrix.
[0050] FIG. 3 illustrates an example stator vane 60. In this
example, the stator vane 60 is a singlet, meaning the stator vane
60 includes only one airfoil section 78. This disclosure is not
limited to singlets, it may be doublets, triplets etc., however.
Further, while the example stator vane 60 is shown in FIG. 2 in the
high pressure turbine section 54, one would understand that this
disclosure can be used in other sections of the engine 20 such as
the mid-turbine frame 57.
[0051] In this example, the airfoil section 78 is connected to both
an inner platform 74 and an outer platform 76. The inner and outer
platforms 74, 76 each respectively include a fore end 74F, 76F, an
aft end 74A, 76A, and circumferential ends 74C, 76C (relative to
the circumferential direction R in FIG. 4). As one would
appreciate, the terms "fore," "aft," and "circumferential" are made
with the normal operational attitude of the stator vane 60. In this
example, flanges 90 are provided on opposite circumferential ends
74C, 76C of both the inner platform 74 and the outer platform
76.
[0052] Each of the flanges 90 is provided with an inner face 901,
an outer face 90F, and a plurality of openings 92 therein. The
inner and outer faces 901, 90F extend radially in the Z direction
and are generally perpendicular to the respective inner and outer
platform 74, 76. Further, each of the openings 92 are configured to
receive fasteners 94 (illustrated schematically in FIGS. 5A-5B) to
fasten adjacent stator vanes 60 together, and to fasten the stator
vanes 60 to a support structure, such as the engine static
structure 36 (FIG. 4).
[0053] With continued reference to FIG. 3, flanges 90 are provided
such that their length dimension L runs in a direction generally
parallel to the engine central longitudinal axis X, and generally
perpendicular to the radial direction Z along the circumferential
ends 74C, 76C of the platforms 74, 76.
[0054] In one example, the stator vane 60 is made of a CMC
material, and the flanges 90 are formed integrally with the stator
vane 60 during a molding process. Because the stator vanes 60 are
formed of a CMC material, they are suitable for use in the high
pressure turbine section 54, however this disclosure is not limited
to a particular section of the gas turbine engine 20.
[0055] FIG. 4 illustrates an example arrangement of a stationary
stage of the gas turbine engine 20, here the first stage array 54a.
In the example, stator vanes 60 are fastened together with
adjacent, similar stator vanes 60 (e.g., the stator vanes 60 each
include platforms 74, 76 configured for circumferential attachment,
as illustrated) such that outer faces 90F of the flanges 90
directly abut one another in a circumferential direction R,
relative to the engine central longitudinal axis X, such that the
outer faces 90F directly abut one another. This arrangement
provides the first stage array 54a with a relatively high hoop
stiffness in the circumferential direction R, which reduces
deflection of the inner and outer platforms 74, 76 and maintains
the overall stability of the first stage array 54a.
[0056] The adjacent stator vanes 60 can further be configured to
carry a seal therebetween, for example, at a point between adjacent
stator vanes 60 (illustrated at 100) to prevent leakage between the
adjacent stator vanes 60.
[0057] FIGS. 5A-5B schematically illustrate examples for fastening
adjacent stator vanes 60 together. In particular, FIGS. 5A-5B
illustrate examples where fasteners 94, which could be nuts and
bolts, are used to fasten the adjacent stator vanes 60 together. In
situations where the vanes and bolts are different materials, the
fasteners 94 can include beveled, or spring, washers to maintain a
preload while accommodating differences in thermal growth. The
fasteners 94 further tie the stator vanes 60 to couplings 95A-95B
such that the adjacent stator vanes are tied together. The
couplings 95A-95B essentially can act as washers and further ensure
a close connection between adjacent stator vanes 60. The outer
coupling 95A is connected to an engine static structure 36 in one
example.
[0058] With further reference to FIGS. 5A-5B, and also with
reference to FIG. 3, the fasteners 94 are positioned within the
openings 92 of adjacent flanges 90 such that the fasteners 94 are
outside the gas flowpath. Thus, exposure of the fasteners 94 to
exhaust gases is reduced.
[0059] With continued reference to FIGS. 5A-5B, resilient members
99 are positioned against the outer platform 76 (FIG. 5A) and
optionally against the inner platform (FIG. 5B) as well. The
resilient members 99 further ensure an adequate connection between
the adjacent stator vanes 60 by providing additional support to the
flanges 90, which aids in taking the load off of the fasteners 94
and further reduces bending of the flanges 90, in particular in the
right and left relative to FIG. 5A.
[0060] In an alternate configuration from those shown in FIGS. 5A
and 5B, the flanges 90 adjacent the outer platform 76 are fastened
together while the flanges 90 adjacent the inner platform 74 are
not fastened together. This allows the inner platforms 74 to move
freely in a radial direction, and thus can prevent fighting between
the adjacent stator vanes 60, and reduces stresses experienced
thereby.
[0061] Although the different examples have the specific components
shown in the illustrations, embodiments of this disclosure are not
limited to those particular combinations. It is possible to use
some of the components or features from one of the examples in
combination with features or components from another one of the
examples.
[0062] One of ordinary skill in this art would understand that the
above-described embodiments are exemplary and non-limiting. That
is, modifications of this disclosure would come within the scope of
the claims. Accordingly, the following claims should be studied to
determine their true scope and content.
* * * * *