Note: Descriptions are shown in the official language in which they were submitted.
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LOW-DUCTILITY TURBINE SHROUD
CROSS-REFERENCE TO RELATED APPLCIATIONS
[0001] This application is a Continuation-In-Part of Application Serial No.
13/327,349,
filed December 15, 2011, which is currently pending.
BACKGROUND OF THE INVENTION
[0002] This invention relates generally to gas turbine engines, and more
particularly to
shrouds made of a low-ductility material in the turbine sections of such
engines.
[0003] A typical gas turbine engine includes a turbomachinery core having a
high
pressure compressor, a combustor, and a high pressure turbine in serial flow
relationship. The
core is operable in a known manner to generate a primary gas flow. The high
pressure turbine
(also referred to as a gas generator turbine) includes one or more rotors
which extract energy
from the primary gas flow. Each rotor comprises an annular array of blades or
buckets carried
by a rotating disk. The flowpath through the rotor is defined in part by a
shroud, which is a
stationary structure which circumscribes the tips of the blades or buckets.
These components
operate in an extremely high temperature environment, and must be cooled by
air flow to
ensure adequate service life. Typically, the air used for cooling is extracted
(bled) from the
compressor. Bleed air usage negatively impacts specific fuel consumption
("SFC") and
should generally be minimized.
[0004] It has been proposed to replace metallic shroud structures with
materials having
better high-temperature capabilities, such as ceramic matrix composites
(CMCs). These
materials have unique mechanical properties that must be considered during
design and
application of an article such as a shroud segment. For example, CMC materials
have
relatively low tensile ductility or low strain to failure when compared with
metallic materials.
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Also, CMCs have a coefficient of thermal expansion ("CTE") in the range of
about 1.5-5
microinch/inch/degree F, significantly different from commercial metal alloys
used as
supports for metallic shrouds. Such metal alloys typically have a CTE in the
range of about
7-10 microinch/inch/degree F.
[0005] CMC materials are comprised of a laminate of a matrix material and
reinforcing
fibers and are orthotropic to at least some degree. The matrix, or non-primary
fiber direction,
herein referred to as interlaminar, is typically weaker (i.e. 1/10 or less)
than the fiber
direction of a composite material system and can be the limiting design
factor.
[0006] Shroud structures are subject to interlaminar tensile stress
imparted at the
junctions between their walls, which must be carried in the weaker matrix
material. These
interlaminar tensile stresses can be the limiting stress location in the
shroud design.
[0007] Accordingly, there is a need for a composite shroud structure with
reduced
interlaminar stresses.
BRIEF DESCRIPTION OF THE INVENTION
[0008] This need is addressed by the present invention, which provides a
shroud
segment configured so as to minimize interlaminar stresses therein.
[0009] According to one aspect of the invention, a shroud segment is
provided for a gas
turbine engine, the shroud segment constructed from a composite material
including
reinforcing fibers embedded in a matrix, and having a cross-sectional shape
defined by
opposed forward and aft walls, and opposed inner and outer walls, the walls
extending
between opposed first and second end faces, wherein the inner wall defines an
arcuate inner
flowpath surface; and wherein a compound fillet is disposed at a junction
between first and
second ones of the walls, the compound fillet including first and second
portions, the second
portion having a concave curvature extending into the first one of the walls.
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[0010] According to another aspect of the invention, a shroud apparatus for
a gas turbine
engine includes: an annular metallic hanger; a shroud segment disposed inboard
of the
hanger, the shroud segment constructed from a composite material including
reinforcing
fibers embedded in a matrix, and having a cross-sectional shape defined by
opposed forward
and aft walls, and opposed inner and outer walls, the walls extending between
opposed first
and second end faces, wherein the inner wall defines an arcuate inner flowpath
surface; and
wherein a compound fillet is disposed at a junction between first and second
ones of the
walls, the compound fillet including first and second portions, the second
portion having a
concave curvature extending into the first one of the walls; and a retainer
mechanically
coupled to the hanger which engages the shroud segment to retain the shroud
segment to the
hanger while permitting movement of the shroud segment in a radial direction.
BRIEF DESCRIPTION OF THE DRAWINGS
[0011] The invention may be best understood by reference to the following
description
taken in conjunction with the accompanying drawing figures in which:
[0012] FIG. 1 is a schematic cross-sectional view of a portion of a turbine
section of a
gas turbine engine, incorporating a shroud mounting apparatus constructed in
accordance
with an aspect of the present invention;
[0013] FIG. 2 is a schematic perspective view of a shroud segment seen in
FIG. 1;
[0014] FIG. 3 is a bottom view of the shroud segment of FIG. 2;
[0015] FIG. 4 is an enlarged view of a portion of FIG. 3;
[0016] FIG. 5 is a sectional front elevation view of a portion of the
turbine section
shown in FIG. 1;
[0017] FIG. 6 is a sectional view of a portion of a shroud segment shown in
FIG. 1;
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[0018] FIG. 7 is a sectional view of a portion of an alternative shroud
segment shown in
FIG. 1; and
[0019] FIG. 8 is a sectional view of a portion of the shroud segment shown
in FIG. 7.
DETAILED DESCRIPTION OF THE INVENTION
[0020] Referring to the drawings wherein identical reference numerals
denote the same
elements throughout the various views, FIG. 1 depicts a small portion of a
turbine, which is
part of a gas turbine engine of a known type. The function of the turbine is
to extract energy
from high-temperature, pressurized combustion gases from an upstream combustor
(not
shown) and to convert the energy to mechanical work, in a known manner. The
turbine drives
an upstream compressor (not shown) through a shaft so as to supply pressurized
air to the
combustor.
[0021] The principles described herein are equally applicable to turbofan,
turbojet and
turboshaft engines, as well as turbine engines used for other vehicles or in
stationary
applications. Furthermore, while a turbine shroud is used as an example, the
principles of the
present invention are applicable to any low-ductility flowpath component which
is at least
partially exposed to a primary combustion gas flowpath of a gas turbine
engine.
[0022] The turbine includes a stationary nozzle 10. It may be of unitary or
built-up
construction and includes a plurality of airfoil-shaped stationary turbine
vanes 12
circumscribed by an annular outer band 14. The outer band 14 defines the outer
radial
boundary of the gas flow through the turbine nozzle 10. It may be a continuous
annular
element or it may be segmented.
[0023] Downstream of the nozzle 10, there is a rotor disk (not shown) that
rotates about
a centerline axis of the engine and carries an array of airfoil-shaped turbine
blades 16. A
shroud comprising a plurality of arcuate shroud segments 18 is arranged so as
to encircle and
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closely surround the turbine blades 16 and thereby define the outer radial
flowpath boundary
for the hot gas stream flowing through the turbine blades 16.
[0024] Downstream of the turbine blades 16, there is a downstream
stationary nozzle 17.
It may be of unitary or built-up construction and includes a plurality of
airfoil-shaped
stationary turbine vanes 19 circumscribed by an annular outer band 21. The
outer band 21
defines the outer radial boundary of the gas flow through the turbine nozzle
17. It may be a
continuous annular element or it may be segmented.
[0025] As seen in FIG. 2, each shroud segment 18 has a generally hollow
cross-sectional
shape defined by opposed inner and outer walls 20 and 22, and forward and aft
walls 24 and
26. Radiused, sharp, or square-edged transitions may be used at the
intersections of the walls.
A shroud cavity 28 is defined within the walls 20, 22, 24, and 26. A
transition wall 29
extends at an angle between the forward wall 24 and the outer wall 22, and
lies at an acute
angle to a central longitudinal axis of the engine when viewed in cross-
section. An axially-
elongated mounting slot 27 passes through the outer wall 22, the transition
wall 29, and the
forward wall 24.The inner wall 20 defines an arcuate radially inner flowpath
surface 30. The
inner wall 20 extends axially forward past the forward wall 24 to define a
forward flange or
overhang 32 and it also extends axially aft past the aft wall 26 to define an
aft flange or
overhang 34. The flowpath surface 30 follows a circular arc in elevation view
(e.g. forward
looking aft or vice-versa).
[0026] The shroud segments 18 are constructed from a ceramic matrix
composite (CMC)
material of a known type. Generally, commercially available CMC materials
include a
ceramic type fiber for example SiC, forms of which are coated with a compliant
material
such as Boron Nitride (BN). The fibers are carried in a ceramic type matrix,
one form of
which is Silicon Carbide (SiC). Typically, CMC type materials have a room
temperature
tensile ductility of no greater than about 1%, herein used to define and mean
a low tensile
ductility material. Generally CMC type materials have a room temperature
tensile ductility in
the range of about 0.4 to about 0.7%. This is compared with metals having a
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temperature tensile ductility of at least about 5%, for example in the range
of about 5 to about
15%. The shroud segments 18 could also be constructed from other low-
ductility, high-
temperature-capable materials.
[0027] CMC materials are orthotropic to at least some degree, i.e. the
material's tensile
strength in the direction parallel to the length of the fibers (the "fiber
direction") is stronger
than the tensile strength in the perpendicular direction (the "matrix",
"interlaminar", or
"secondary" or "tertiary" fiber direction). Physical properties such as
modulus and Poisson's
ratio also differ between the fiber and matrix directions.
[0028] The flowpath surface 30 of the shroud segment 18 may incorporate a
layer of
environmental barrier coating ("EBC"), which may be an abradable material,
and/or a rub-
tolerant material of a known type suitable for use with CMC materials. This
layer is
sometimes referred to as a "rub coat", designated at 38. As used herein, the
term "abradable"
implies that the rub coat 38 is capable of being abraded, ground, or eroded
away during
contact with the tips of the turbine blades 16 as they turn inside the shroud
segments 18 at
high speed, with little or no resulting damage to the turbine blade tips. This
abradable
property may be a result of the material composition of the rub coat 38, by
its physical
configuration, or by some combination thereof The rub coat 38 may comprise a
ceramic
layer, such as yttria stabilized zirconia or barium strontium alumino
silicate. Exemplary
compositions and methods suitable for making the rub coat 38 are described in
U.S. Pat. No.
7,749,565 (Johnson et al.), which is incorporated herein by reference.
[0029] FIGS. 3 and 4 depict the rub coat 38 in more detail. In the
illustrated example, the
rub coat 38 is patterned. The pattern enhances abradability of the rub coat by
decreasing the
surface area exposed to contact with the tips of the turbine blades 16.
Specifically, the rub
coat 38 has a plurality of side-by-side grooves 39 formed therein. The
presence of the
grooves 39 gives the surface a shape comprising alternate peaks 41 and valleys
43. The
grooves 39 run generally in a fore-to-aft direction, and each groove 39 has a
forward end 45,
a central portion 47, and an aft end 49. In plan view, the grooves 39 may be
curved. For
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example, as shown in FIG. 3, each groove 39 is curved such that its central
portion 47 is
offset in a lateral or tangential direction relative to its forward and aft
ends 45 and 49.
[0030] The shroud segments 18 include opposed end faces 42 (also commonly
referred
to as "slash" faces). The end faces 42 may lie in a plane parallel to the
centerline axis of the
engine, referred to as a "radial plane", or they may be slightly offset from
the radial plane, or
they may be oriented so that they are at an acute angle to such a radial
plane. When
assembled into a complete ring, end gaps are present between the end faces 42
of adjacent
shroud segments 18. One or more seals (not shown) may be provided at the end
faces 42.
Similar seals are generally known as "spline seals" and take the form of thin
strips of metal or
other suitable material which are inserted in slots in the end faces 42. The
spline seals span
the gaps between shroud segments 18.
[0031] FIG. 6 illustrates the interior construction of the shroud segment
18 in more
detail. There is a concave fillet 19 present between the inner wall 22 and the
aft wall 26. This
fillet 19 is representative of the junctions present at each of the four
intersections where two
of the four walls meet each other. In operation, this type of configuration
can experience a
peak interlaminar tensile stress below the surface of the material, near the
location of the
fillet 19, which must be carried in the weaker matrix material. This can be
the limiting stress
location in the design of the shroud segment 18.
[0032] FIG. 7 illustrates an alternative shroud segment 118. The basic
configuration is
similar to that of the shroud segment 18, but the shroud segment 118 is
configured to reduce
the interlaminar stresses in the composite material. It has a generally hollow
cross-sectional
shape defined by opposed inner and outer walls 120 and 122, and forward and
aft walls 124
and 126. A shroud cavity 128 is defined within the walls 120, 122, 124, and
126. A
compound fillet 119 is present between the inner wall 122 and the aft wall
126. This fillet
119 is representative of the junctions present at each of the four
intersections where two of
the four walls meet each other.
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[0033] As best seen in FIG. 8, the compound fillet 119 includes a first
portion 119A
which has a surface disposed at an acute angle to the interior surface of the
aft wall 126 and
the interior surface of the inner wall 120. The surface of the first portion
119A may be
generally flat. The first portion 119A represents an addition of material
relative to the
nominal thickness of the aft wall 126, as seen by the location of the dashed
line 130. The
compound fillet 119 also includes a second portion 119B which is concave-
curved surface
having a radius R. A first end 132 of the second portion 119B meets the first
portion 119A,
and a second end 134 of the second portion 119B meets and transitions to the
interior surface
of the inner wall 120. The second portion 119B represents a subtraction of
material relative
to the nominal thickness of the aft wall 126, as seen by the location of the
dashed line 136.
The compound fillet 119, particularly the second portion 119B, may be
considered an
"undercut" or "thinning" preceding or adjacent to a concentrated
interlaminated stress region.
[0034] At the junction of the first portion 119A and the interior surface
of the aft wall
126, there is a first transition surface 138, which is illustrated as a smooth
concave curve.
Other configurations which could produce similar results include straight
lines or spline
shapes.
[0035] A second transition portion 140 is disposed at the junction of the
second portion
119B and the interior surface of the inner wall 120, which is illustrated as a
smooth convex
curve. Other configurations which could produce similar results include
straight lines or
spline shapes.
[0036] The profile of the compound fillet 119 is shaped so as to be
compatible with
composite materials. The reinforcing fibers within the component generally
follow the
contours of (i.e. are parallel to) the bounding surfaces of the interior wall
120, the compound
fillet 119, and the aft wall 126. These surfaces are contoured such that the
fibers will not
buckle or wrinkle where outward cusps are located. While the profile of the
compound fillet
119 has been illustrated in an exemplary two-dimensional sectional view, it is
noted that the
actual shape may be different at different sections.
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[0037] In the illustrated example, the thickness of the inner wall 120 is
at a minimum at
the location of the second portion 119B of the compound fillet 119. The exact
shapes and
dimensions of the compound fillet 119 may be altered to suit a particular
application and the
specific composite material used.
[0038] The compound fillet 119 has been illustrated disposed between the
aft wall 126
and the forward wall 120. It is noted that the same or similar configuration
may be
implemented at the junctions between any or all of the walls 120, 122, 124,
and 126.
[0039] The shroud segments 18 are mounted to a stationary metallic engine
structure,
shown in FIG. 1. In this example the stationary structure is part of a turbine
case 44. The ring
of shroud segments 18 is mounted to an array of arcuate shroud hangers 46 by
way of an
array of retainers 48 and bolts 50.
[0040] As best seen in FIGS. 1 and 5, each hanger 46 includes an annular
body 52 which
extends in a generally axial direction. The body 52 is angled such that its
forward end is
radially inboard of its aft end. It is penetrated at intervals by radially-
aligned bolt holes 54.
An annular forward outer leg 56 is disposed at the forward end of the body 52.
It extends in a
generally radial direction outboard of the body 52, and includes a forward
hook 58 which
extends axially aft. An annular aft outer leg 60 is disposed at the aft end of
the body 52. It
extends in a generally radial direction outboard of the body 52, and includes
an aft hook 62
which extends axially aft. An annular forward inner leg 64 is disposed at the
forward end of
the body 52. It extends in a generally radial direction inboard of the body
52, and includes an
aft-facing, annular forward bearing surface 66. An annular aft inner leg 68 is
disposed at the
aft end of the body 52. It extends in a generally radial direction inboard of
the body 52, and
includes a forward-facing, annular aft bearing surface 70. As will be
explained in more detail
below, the aft inner leg 68 is configured to function as a spring element. The
body 52 has one
or more coolant feed passages 71 formed therein which serve to receive coolant
from a
source within the engine (such as compressor bleed air) and route the coolant
to the inboard
side of the body 52.
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[0041] The
hangers 46 are installed into the turbine case 44 as follows. The forward
hook 58 is received by an axially-forward facing forward rail 72 of the case
44. The aft hook
62 is received by an axially-forward facing aft rail 74 of the case 44. An
anti-rotation pin 76
or other similar anti-rotation feature is received in the forward rail 72 and
extends into a
mating slot (not shown) in the forward hook 58.
[0042] The
construction of the retainers 48 is shown in more detail in FIG. 5. Each
retainer 48 has a central portion 78 with two laterally-extending arms 80. The
distal end of
each arm 80 includes a concave-curved contact pad 82 which protrudes radially
outward
relative to the remainder of the arm 80. The central portion 78 is raised
above the arms 80 in
the radial direction and defines a clamping surface 84. A radially-aligned
bore 86 extends
through the central portion 78. A generally tubular insert 88 is swaged or
otherwise secured
to the bore 86 and includes a threaded fastener hole. Optionally, the bore 86
could be
threaded and the insert 88 eliminated.
[0043] The
retainer 48 is positioned in the shroud cavity 28 with the central portion 78
and the clamping surface 84 exposed through the mounting hole 27 in the outer
wall 22. The
retainer 48 is clamped against a boss 90 of the hanger 46 by the bolt 50 or
other suitable
fastener, and a spring 92 is clamped between the boss 90 and the clamping
surface. Each
spring 92 includes a center section with a mounting hole, and opposed
laterally-extending
arms 94.
[0044] The
relative dimensions of the boss 90, the retainer 48, and the shroud segment
18 are selected such that the retainers 48 limit the inboard movement of the
shroud segments
18, but do not clamp the shroud segments 18 against the hanger 46 in the
radial direction. In
other words, the retainers 48 permit a definite clearance for movement in the
radially
outboard direction. In operation, the prevailing gas pressure load in the
secondary flowpath
urges the shroud segment 18 radially inboard against the retainer 48, while
the retainer 48
deflects a small amount.
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[0045] The springs 92 function to hold the shroud segments 18 radially
inboard against
the retainers 48 during assembly and for an initial grinding process to
circularize the ring of
shroud segments 18. However, the springs 92 are sized such that they do not
exert a
substantial clamping load on the shroud segments 18.
[0046] In the axial direction, the aft inner leg 68 of the hanger 46 acts
as a large
cantilevered spring to counteract air pressure loads in operation. This spring
action urges the
forward wall 24 of the shroud segment 18 against the forward bearing surface
66 of the
forward inner leg 64, resulting in a positive seal between the metallic hanger
46 and the
CMC shroud segments, thereby decreasing cooling flow leakage.
[0047] In the installed condition, the forward and aft overhangs 32 and 34
are disposed
in axially close proximity or in axially overlapping relationship with the
components forward
and aft of the shroud segment 18. In the illustrated example, there is an
overlapping
configuration between the aft overhang 34 and the aft nozzle band 21, while
the forward
overhang 32 lies in close proximity to the forward outer band 14. This
configuration
minimizes leakage between the components and discourages hot gas ingestion
from the
primary flowpath to the secondary flowpath.
[0048] As noted above, the mounting slot 27 passes through the outer wall
22, the
transition wall 29, and the forward wall 24. The shroud segments 18 thus
incorporate a
substantial amount of open area. There is not an air seal present between the
perimeter of the
mounting slot 27 and the hanger 46, and the shroud segments 18 do not, in and
of
themselves, function as plenums. Rather, the shroud segments 18 form a plenum
in
cooperation with the hangers 46, indicated generally at "P" in FIG. 1.
Specifically, an annular
sealing contact is present between the forward bearing surface 66 and the
forward wall 24 of
the shroud segment 18. Also, an annular sealing contact is present between the
aft bearing
surface 70 and the aft wall 26 of the shroud segment 18. The sealing contact
is ensured by the
spring action of the aft inner leg 68 as described above. The shroud segments
18 may be
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considered to be the "inner portion" of the plenum and the hangers 46 may be
considered to
be the "outer portion" thereof
[0049] A hollow metallic impingement baffle 96 is disposed inside each
shroud segment
18. The impingement baffle 96 fits closely to the retainer 48. The inboard
wall of the
impingement baffle has a number of impingement holes 98 formed therein, which
direct
coolant at the segment 18. The interior of the impingement baffle 96
communicates with the
coolant feed passage 71 through a transfer passage 73 formed in the retainer
48.
[0050] In operation, air flows through passage 71, transfer passage 73,
baffle 96,
impingement holes 98, and pressurizes the plenum P. Spent cooling air from the
plenum P
exits through purge holes 100 formed in the forward wall 24 of the shroud
segment 18.
[0051] The shroud mounting apparatus described above is effective to mount
a low-
ductility shroud in a turbine engine without applying clamping loads directly
thereto, and has
several advantages compared to the prior art.
[0052] In particular, the tapered edge (or wedge) shape on the forward side
of the shroud
allows the shroud mounting system to carry loads from forward of the shroud
segments 18 to
the turbine case 44 without transmitting directly through the shroud segments
18. By
redirecting the load around the shroud segments 18, the stress in the shroud
segments 18
remains relatively low.
[0053] Furthermore, the overhangs 32 and 34 allow the shroud segments 18 to
protect
the supporting structure close to the flowpath while discouraging hot gas
ingestion through
the use of overlaps between the shroud segments 18 and the axially adjacent
nozzles. This
overlapping configuration requires less cooling flow to purge the shroud-to-
nozzle cavities,
thereby improving overall engine performance. As the shroud material has
better high
temperature capability and lower stress than the adjacent nozzles, the use of
the overhangs 32
and 34 provides an overall turbine life improvement.
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[0054] Finally, the incorporation of the compound fillet 119 allows the
interlaminar
stress at the shroud segment wall intersections to be distributed over a
larger area, thus
reducing the peak interlaminar tensile stress value. Analysis has shown that
the configuration
described above can lower the peak interlaminar tensile stress by a
significant amount, for
example about 50% as compared to the configuration without the compound
fillet, without
significant changes to the primary in-plane (or fiber direction) stress.
[0055] The foregoing has described a turbine shroud apparatus for a gas
turbine engine.
While specific embodiments of the present invention have been described, it
will be apparent
to those skilled in the art that various modifications thereto can be made
without departing
from the spirit and scope of the invention. Accordingly, the foregoing
description of the
preferred embodiment of the invention and the best mode for practicing the
invention are
provided for the purpose of illustration only and not for the purpose of
limitation.
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