U.S. patent application number 13/723941 was filed with the patent office on 2014-09-11 for manufacture of full ring strut vane pack.
This patent application is currently assigned to UNITED TECHNOLOGIES CORPORATION. The applicant listed for this patent is United Technologies Corporation. Invention is credited to Shelton O. Duelm, Michael G. McCaffrey.
Application Number | 20140255174 13/723941 |
Document ID | / |
Family ID | 50979196 |
Filed Date | 2014-09-11 |
United States Patent
Application |
20140255174 |
Kind Code |
A1 |
Duelm; Shelton O. ; et
al. |
September 11, 2014 |
MANUFACTURE OF FULL RING STRUT VANE PACK
Abstract
A vane assembly for a gas turbine engine is disclosed and
includes a plurality of individual vane segments forming an annular
ring. Each of the plurality of vane segments is formed of a ceramic
matrix composite including at least one airfoil extending between
an inner platform and an outer platform. An outer wrap is disposed
about an outer periphery of the annular ring and over an interface
between outer platforms of adjacent ones of the plurality of vane
segments.
Inventors: |
Duelm; Shelton O.;
(Wethersfield, CT) ; McCaffrey; Michael G.;
(Windsor, CT) |
|
Applicant: |
Name |
City |
State |
Country |
Type |
United Technologies Corporation; |
|
|
US |
|
|
Assignee: |
UNITED TECHNOLOGIES
CORPORATION
Hartford
CT
|
Family ID: |
50979196 |
Appl. No.: |
13/723941 |
Filed: |
December 21, 2012 |
Current U.S.
Class: |
415/200 ;
156/185; 415/210.1 |
Current CPC
Class: |
F05D 2300/6033 20130101;
F01D 5/284 20130101; F01D 9/041 20130101; Y02T 50/60 20130101; Y02T
50/672 20130101; F01D 25/162 20130101; F05D 2230/23 20130101; Y02T
50/673 20130101 |
Class at
Publication: |
415/200 ;
415/210.1; 156/185 |
International
Class: |
F01D 5/28 20060101
F01D005/28; F01D 9/04 20060101 F01D009/04 |
Claims
1. A vane assembly for a gas turbine engine comprising: a plurality
of individual vane segments forming an annular ring with each of
the plurality of vane segments formed of a ceramic matrix composite
including at least one airfoil extending between an inner platform
and an outer platform; and an wrap disposed about one of an outer
periphery and an inner periphery of the annular ring and over an
interface between a corresponding inner and outer platform of
adjacent ones of the plurality of vane segments.
2. The vane assembly as recited in claim 1, wherein the wrap
comprises an outer wrap disposed about the outer periphery and
covering the interface between outer platforms of adjacent ones of
the plurality of vane segments.
3. The vane assembly as recited in claim 1, wherein the wrap
comprises an inner wrap disposed about an inner periphery of the
annular ring over an interface between inner platforms of adjacent
ones of the plurality of vane segments.
4. The vane assembly as recited in claim 1, wherein the wrap
comprise a ceramic matrix composite material.
5. The vane assembly as recited in claim 1, wherein the wrap
includes a silicon carbide monofilament.
6. The vane assembly as recited in claim 1, wherein the wrap
comprises a plurality of plies applied to the plurality of vane
segments.
7. The vane assembly as recited in claim 1, wherein each of the
plurality of vane segments include a cavity extending through the
outer platform, airfoils and inner platform.
8. The vane assembly as recited in claim 1, wherein the plurality
of vane segments are at least partially cured prior to installation
of the wrap.
9. A mid-turbine frame for a gas turbine engine comprising: a
plurality of individual vane segments forming an annular ring with
each of the plurality of vane segments formed of a ceramic matrix
composite including at least one airfoil extending between an inner
platform and an outer platform; an outer wrap disposed about an
outer periphery of the annular ring over an interface between outer
platforms of adjacent ones of the plurality of vane segments; and a
support tie-rod extending through at least one of the plurality of
vane segments and attached at a first end to an outer case
structure and on a second end to an inner case structure.
10. The mid-turbine frame as recited in claim 9, including an inner
wrap disposed about an inner periphery of the annular ring over an
interface between inner platforms of adjacent ones of the plurality
of vane segments.
11. The vane assembly as recited in claim 10, wherein the outer
wrap and the inner wrap comprise a ceramic matrix composite
material.
12. The vane assembly as recited in claim 11, wherein at least one
of the outer wrap and inner wrap includes a silicon carbide
monofilament.
13. The vane assembly as recited in claim 10, wherein at least one
of the outer wrap and inner wrap comprise a plurality of plies
applied to the plurality of vane segments.
14. A method of assembling a full ring vane pack comprising:
forming a plurality of vane segments from a ceramic matrix
composite material including at least one airfoils extending
between an outer platform and an inner platform; assembling an
annular ring with the plurality of vane segments by abutting
corresponding outer platforms and inner platforms; and wrapping an
outer periphery of the annular ring with a ceramic matrix composite
material over an interface between outer platforms of adjacent ones
of the plurality of vane segments.
15. The method as recited in claim 14, including wrapping an inner
periphery of the annular ring with a ceramic matrix composite
material over an interface between inner platforms of adjacent ones
of the plurality of vane segments.
16. The method as recited in claim 14, including curing the wrapped
ceramic matrix composite material around the plurality of vane
segments.
17. The method as recited in claim 14, wherein each of the
plurality of vane segments are fully cured prior to assembly into
the annular ring.
18. The method as recited in claim 14, wherein each of the
plurality of vane segments are not fully cured when assembled into
the annular ring and both the outer ceramic matrix composite
wrapping and plurality of vane segments are finish cured together.
Description
BACKGROUND
[0001] A gas turbine engine typically includes a fan section, a
compressor section, a combustor section and a turbine section. Air
entering the compressor section is compressed and delivered into
the combustion section where it is mixed with fuel and ignited to
generate a high-speed exhaust gas flow. The high-speed exhaust gas
flow expands through the turbine section to drive the compressor
and the fan section. The compressor section typically includes low
and high pressure compressors, and the turbine section includes low
and high pressure turbines.
[0002] Blades and vanes within the turbine section encounter
elevated temperatures that require either provisions for cooling or
the use of temperature compatible materials. Moreover, it is
desired to limit leakage of hot gases between adjacent vanes in the
turbine section and therefore sealing features that are also
compatible with the elevated temperatures within the gas path are
utilized.
[0003] Accordingly, it is desirable to design and develop turbine
elements such as vanes that balance performance and expense to
provide a desired performance and cost effectiveness. Although
current gas turbine engines are increasingly more efficient,
turbine engine manufacturers continue to seek further improvements
to engine performance including improvements to thermal, transfer
and propulsive efficiencies.
SUMMARY
[0004] A vane assembly for a gas turbine engine according to an
exemplary embodiment of this disclosure, among other possible
things includes a plurality of individual vane segments forming an
annular ring with each of the plurality of vane segments formed of
a ceramic matrix composite including at least one airfoil extending
between an inner platform and an outer platform, and a wrap
disposed about one of an outer periphery and an inner periphery of
the annular ring and over an interface between a corresponding
inner and outer platform of adjacent ones of the plurality of vane
segments.
[0005] In a further embodiment of the foregoing vane assembly, the
wrap includes an outer wrap disposed about the outer periphery and
covering the interface between outer platforms of adjacent ones of
the plurality of vane segments.
[0006] In a further embodiment of any of the foregoing vane
assemblies, the wrap includes an inner wrap disposed about an inner
periphery of the annular ring over an interface between inner
platforms of adjacent ones of the plurality of vane segments.
[0007] In a further embodiment of any of the foregoing vane
assemblies, the wrap includes a ceramic matrix composite
material.
[0008] In a further embodiment of any of the foregoing vane
assemblies, the wrap includes a silicon carbide monofilament.
[0009] In a further embodiment of any of the foregoing vane
assemblies, the wrap includes a plurality of plies applied to the
plurality of vane segments.
[0010] In a further embodiment of any of the foregoing vane
assemblies, each of the plurality of vane segments include a cavity
extending through the outer platform, airfoils and inner
platform.
[0011] In a further embodiment of any of the foregoing vane
assemblies, the plurality of vane segments are at least partially
cured prior to installation of the wrap.
[0012] A mid-turbine frame for a gas turbine engine according to an
exemplary embodiment of this disclosure, among other possible
things includes a plurality of individual vane segments forming an
annular ring with each of the plurality of vane segments formed of
a ceramic matrix composite including at least one airfoil extending
between an inner platform and an outer platform. An outer wrap is
disposed about an outer periphery of the annular ring over an
interface between outer platforms of adjacent ones of the plurality
of vane segments. A support tie-rod extends through at least one of
the plurality of vane segments and is attached at a first end to an
outer case structure and on a second end to an inner case
structure.
[0013] In a further embodiment of the foregoing mid-turbine frame,
includes an inner wrap disposed about an inner periphery of the
annular ring over an interface between inner platforms of adjacent
ones of the plurality of vane segments.
[0014] In a further embodiment of any of the foregoing vane
assemblies, the outer wrap and the inner wrap include a ceramic
matrix composite material.
[0015] In a further embodiment of any of the foregoing vane
assemblies, at least one of the outer wrap and inner wrap includes
a silicon carbide monofilament.
[0016] In a further embodiment of any of the foregoing vane
assemblies, at least one of the outer wrap and inner wrap include a
plurality of plies applied to the plurality of vane segments.
[0017] A method of assembling a full ring vane pack according to an
exemplary embodiment of this disclosure, among other possible
things includes forming a plurality of vane segments from a ceramic
matrix composite material including at least one airfoils extending
between an outer platform and an inner platform, assembling an
annular ring with the plurality of vane segments by abutting
corresponding outer platforms and inner platforms, and wrapping an
outer periphery of the annular ring with a ceramic matrix composite
material over an interface between outer platforms of adjacent ones
of the plurality of vane segments.
[0018] In a further embodiment of the foregoing method, includes
wrapping an inner periphery of the annular ring with a ceramic
matrix composite material over an interface between inner platforms
of adjacent ones of the plurality of vane segments.
[0019] In a further embodiment of any of the foregoing methods,
includes curing the wrapped ceramic matrix composite material
around the plurality of vane segments.
[0020] In a further embodiment of any of the foregoing methods,
each of the plurality of vane segments are fully cured prior to
assembly into the annular ring.
[0021] In a further embodiment of any of the foregoing methods,
each of the plurality of vane segments are not fully cured when
assembled into the annular ring and both the outer ceramic matrix
composite wrapping and plurality of vane segments are finish cured
together.
[0022] 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.
[0023] These and other features disclosed herein can be best
understood from the following specification and drawings, the
following of which is a brief description.
BRIEF DESCRIPTION OF THE DRAWINGS
[0024] FIG. 1 is a schematic view of an example gas turbine
engine.
[0025] FIG. 2 is a section looking from the rear to the front
through a center of an example mid turbine frame.
[0026] FIG. 3 is a close up cross-section of the gas turbine
engine.
[0027] FIG. 4 is a side view of an example vane pack assembly.
[0028] FIG. 5 is an enlarged view of the example vane pack
assembly.
[0029] FIG. 6 is a perspective view of an example vane segment.
DETAILED DESCRIPTION
[0030] 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.
[0031] 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.
[0032] 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 A 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.
[0033] 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 A.
[0034] 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.
[0035] 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.
[0036] A mid-turbine frame 58 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 58 further supports
bearing systems 38 in the turbine section 28 as well as setting
airflow entering the low pressure turbine 46.
[0037] Airflow through the core flow 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 58 includes vanes 60, which are in the core airflow path and
function as an inlet guide vane for the low pressure turbine 46.
Utilizing the vane 60 of the mid-turbine frame 58 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 58. 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.
[0038] 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.
[0039] 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.
[0040] 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.
[0041] "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.
[0042] "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.
[0043] 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.
[0044] Referring to FIGS. 2 and 3 with continued reference to FIG.
1, the example mid turbine frame 58 includes a tie rod 62 that
extends from an outer case 66 through a vane pack 68 to an inner
case 64. The tie rods 62 extend through the vane pack 68 and
specifically through openings within individual vane segments
70.
[0045] The example vane pack 68 includes a plurality of vane
segments that define an annular ring about the engine axis A. The
vane pack 68 defines a gas flow path between the high pressure
turbine 54 and the low pressure turbine 46. Moreover, the mid
turbine frame 58 provides for the support of bearing structures
that support rotation of the high and low shafts 40, 50.
[0046] Referring to FIGS. 4, 5 and 6, the example vane pack
assembly 68 includes ceramic matrix composite vane segments 70 that
includes the airfoil 60 disposed between an inner platform 72 and
an outer platform 74. A plurality of vane segments 70 are arranged
in an annular ring that includes an inner periphery and an outer
periphery. The inner periphery is defined by the inner platform 72
and the outer periphery is defined by the outer platform 74.
[0047] It should be understood that although the disclosed vane
segments 70 include only a single airfoil 60 between the platforms
72,74, referred commonly to as a singlet, other vane segment
configurations that include more than one airfoil 60 between common
platforms 72, 74 commonly known as doublets are also within the
contemplation of this disclosure. Moreover, doublets, triplets or
any number of vane segment configurations are within the
contemplation of this disclosure.
[0048] Each of the individual vane segments 70 are fabricated from
a ceramic matrix composite material as an individual unit and
assembled into the desired annular ring configuration illustrated
in FIG. 4. Once the individual vane segments 70 are arranged in the
annular ring configuration, an outer wrap 76 is wrapped around the
outer periphery defined by the outer platform 74. The outer wrap 76
extends over an interface 80 between abutting individual vane
segments 70.
[0049] In this example, the outer wrap 76 also comprises a ceramic
matrix composite wrap that is wrapped about the entire
circumference of the vane pack assembly 68. The outer wrap 76 may
also include a silicon carbide monofilaments commonly referred to
SCS-6. The use of the silicon carbide monofilaments adds additional
strength and provides a full hoop structure along with the ceramic
matrix composite wrap about the outer platforms to provide the
desired structural rigidity of the vane pack assembly 68.
[0050] An inner wrap 78 is also provided that overlaps interfaces
82 between adjacent inner platforms 72 of the various individual
vane segments 70. The inner wrap 78 may also include the silicon
carbide monofilament material 84 to provide the desired additional
strength and structural rigidity.
[0051] The example vane pack 68 is fabricated by first assembling a
plurality of individually cured vane segments 70. The vane segments
70 may be fully cured or partially cured. Each of the vane segments
70 are individually formed and assembled in the desired annular
ring configuration.
[0052] The annular ring configuration includes the inner ring
defined by the inner platform 72 with the airfoil segments 60
extending from the inner platform 72 to the outer platform 74 that
defines the outer ring. An interface 80 between each of the
adjacent outer platforms 74 is covered by a wrap material 76. In
this example, the wrap material 76 comprises a ceramic matrix
composite fabric material wrapped about the outer periphery to
overlap each of the interfaces 80 between adjacent vane segments
70. The over wrap material 76 is applied in plies to build a layer
of material over the interfaces 80 to further reduce leakage and
provide the desired structural rigidity.
[0053] In one example, the example vane segments 70 are assembled
in the annular ring configuration when in a partially cured
condition. The over wrap material 76 is then wrapped about the
outer periphery also in an uncured condition. The entire assembly
is then finished cured at the same time to provide the desired
finished structure in a fully cured state.
[0054] In a further example, each of the plurality of vane segments
70 is fully cured prior to assembly into the annular ring
configuration. Prior curing and fabrication of each of the
plurality of segments 70 enables the use of structural segments to
form the entire vane pack assembly 68. Prior configurations sought
to fabricate the entire vane pack assembly from a ceramic matrix
composite material all at the same time and incurred substantial
manufacturing expense and provided unpredictable results.
[0055] Accordingly, the example method and vane pack assembly
provides a ceramic matrix composite structure that can be
fabricated from individually cured vane segments to reduce the
manufacturing cost and complexity. The addition of the wrap
material 76 about the outer periphery reduces leakage through
interfaces 80 between adjacent vane segments 70. Moreover, the over
wrap 76 provides the desired structural rigidity and a full hoop
support structure capable of maintaining the desired structure
rigidity of the vane pack assembly 68. The example wraps 76 and 78
are applied in a flexible material sheet form and cured along with
the finished curing of the individual vane segments.
[0056] Accordingly, the example vane pack assembly provides a
ceramic matrix composite assembly that can be utilized throughout
the turbine section and that provides low leakage through the
interfaces between the adjacent vane segments 70. Moreover, the
disclosed assembly technique and method enables more efficient
fabrication of a strut ring vane pack. The example ring structure
as provided by the example vane pack 68 further reduces the amount
of leakage and provides for the separate vanes 70 to be assembled
into a ring to provide the desired flow area and aerodynamic
characteristics.
[0057] Although an example embodiment has been disclosed, a worker
of ordinary skill in this art would recognize that certain
modifications would come within the scope of this disclosure. For
that reason, the following claims should be studied to determine
the scope and content of this disclosure.
* * * * *