Note: Descriptions are shown in the official language in which they were submitted.
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TURBINE ENGINE SHROUD ASSEMBLY
FIELD OF THE INVENTION
[0001] The present invention generally relates to turbine engines and in
particular
relates to a turbine engine shroud assembly.
BACKGROUND OF THE INVENTION
[0002] Turbine engines, and particularly gas or combustion turbine engines,
are rotary
engines that extract energy from a flow of pressurized combusted gases passing
through
the engine onto a multitude of rotating turbine blades.
[0003] The rotating turbine blades can be supported by shrouds that are
interlocked to
form a circumferential casing to the turbine. A Z-shaped interlock is a
typical configuration
choice for a shrouded blade assembly which requires a pre-twist during
manufacturing and
assembly. Eliminating the pre-twist while maintaining an interlock
configuration would be
beneficial for shroud assembly manufacturing.
BRIEF DESCRIPTION OF THE INVENTION
[0004] In one aspect, embodiments of relate to a turbine engine comprising
a rotor
having a plurality of radially extending airfoils spaced circumferentially
about the rotor,
with the airfoils terminating in a tip, and a shroud assembly circumscribing
the airfoils and
comprising a shroud element mounted to each tip and having opposing radial
sides with
first and second interlock elements, wherein the first interlock element of
one shroud
element mates with a second interlock element of a circumferentially adjacent
element to
form a plurality of interlocks between adjacent shroud elements about the
circumference
of the airfoils.
[0005] In another aspect, embodiments relate to an interlocking shroud
assembly for a
turbine engine comprising a plurality of radially extending, circumferentially
spaced
airfoils terminating in a shroud element and having opposing radial sides with
first and
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second interlock elements, which interlock to form interlocks between
circumferentially
adjacent airfoils.
[0006] In yet another aspect, embodiments relate to a method of forming a
shroud about
a plurality of rotating blades in a turbine engines comprising forming an
interlock between
circumferentially adjacent tips of the blades and preloading the interlock.
BRIEF DESCRIPTION OF THE DRAWINGS
[0007] In the drawings:
[0008] Figure 1 is a schematic cross-sectional diagram of a turbine engine
for an
aircraft.
[0009] Figure 2 is an assembled plurality of airfoils.
[0010] Figure 3 is a perspective view of a shroud element.
[0011] Figure 4 is another perspective view of a shroud element.
[0012] Figure 5 is an illustration of a shroud assembly.
[0013] Figure 6 is a cross-sectional view of the shroud assembly of Figure
5.
[0014] Figure 7 is a cross-sectional view of a second embodiment of the
shroud
assembly of Figure 5.
[0015] Figure 8 is a single airfoil assembly.
DESCRIPTION OF EMBODIMENTS OF THE INVENTION
[0016] The described embodiments of the present invention are directed to a
shroud
assembly for an airfoil. For purposes of illustration, the present invention
will be described
with respect to the turbine for an aircraft turbine engine. It will be
understood, however,
that the invention is not so limited and may have general applicability within
an engine,
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including compressors, as well as in non-aircraft applications, such as other
mobile
applications and non-mobile industrial, commercial, and residential
applications.
[0017] As used herein, the term "forward" or "upstream" refers to moving in
a direction
toward the engine inlet, or a component being relatively closer to the engine
inlet as
compared to another component. The term "aft" or "downstream" used in
conjunction with
"forward" or "upstream" refers to a direction toward the rear or outlet of the
engine relative
to the engine centerline.
[0018] Additionally, as used herein, the terms "radial" or "radially" refer
to a
dimension extending between a center longitudinal axis of the engine and an
outer engine
circumference.
[0019] All directional references (e.g., radial, axial, proximal, distal,
upper, lower,
upward, downward, left, right, lateral, front, back, top, bottom, above,
below, vertical,
horizontal, clockwise, counterclockwise, upstream, downstream, aft, etc.) are
only used for
identification purposes to aid the reader's understanding of the present
invention, and do
not create limitations, particularly as to the position, orientation, or use
of the invention.
Connection references (e.g., attached, coupled, connected, and joined) are to
be construed
broadly and can include intermediate members between a collection of elements
and
relative movement between elements unless otherwise indicated. As such,
connection
references do not necessarily infer that two elements are directly connected
and in fixed
relation to one another. The exemplary drawings are for purposes of
illustration only and
the dimensions, positions, order and relative sizes reflected in the drawings
attached hereto
can vary.
[0020] Figure 1 is a schematic cross-sectional diagram of a turbine engine
10 for an
aircraft. The engine 10 has a generally longitudinally extending axis or
centerline 12
extending forward 14 to aft 16. The engine 10 includes, in downstream serial
flow
relationship, a fan section 18 including a fan 20, a compressor section 22
including a
booster or low pressure (LP) compressor 24 and a high pressure (HP) compressor
26, a
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combustion section 28 including a combustor 30, a turbine section 32 including
a HP
turbine 34, and a LP turbine 36, and an exhaust section 38.
[0021] The fan section 18 includes a fan casing 40 surrounding the fan 20.
The fan 20
includes a plurality of fan blades 42 disposed radially about the centerline
12. The HP
compressor 26, the combustor 30, and the HP turbine 34 form a core 44 of the
engine 10,
which generates combustion gases. The core 44 is surrounded by core casing 46,
which
can be coupled with the fan casing 40.
[0022] A HP shaft or spool 48 disposed coaxially about the centerline 12 of
the engine
drivingly connects the HP turbine 34 to the HP compressor 26. A LP shaft or
spool 50,
which is disposed coaxially about the centerline 12 of the engine 10 within
the larger
diameter annular HP spool 48, drivingly connects the LP turbine 36 to the LP
compressor
24 and fan 20.
[0023] The LP compressor 24 and the HP compressor 26 respectively include a
plurality of compressor stages 52, 54, in which a set of compressor blades 56,
58 rotate
relative to a corresponding set of static compressor vanes 60, 62 (also called
a nozzle) to
compress or pressurize the stream of fluid passing through the stage. In a
single compressor
stage 52, 54, multiple compressor blades 56, 58 can be provided in a ring and
can extend
radially outwardly relative to the centerline 12, from a blade platform to a
blade tip, while
the corresponding static compressor vanes 60, 62 are positioned upstream of
and adjacent
to the rotating blades 56, 58. It is noted that the number of blades, vanes,
and compressor
stages shown in Figure 1 were selected for illustrative purposes only, and
that other
numbers are possible.
[0024] The blades 56, 58 for a stage of the compressor can be mounted to a
disk 59,
which is mounted to the corresponding one of the HP and LP spools 48, 50, with
each stage
having its own disk 59, 61. The vanes 60, 62 for a stage of the compressor can
be mounted
to the core casing 46 in a circumferential arrangement.
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[0025] The HP turbine 34 and the LP turbine 36 respectively include a
plurality of
turbine stages 64, 66, in which a set of turbine blades 68, 70 are rotated
relative to a
corresponding set of static turbine vanes 72, 74 (also called a nozzle) to
extract energy from
the stream of fluid passing through the stage. In a single turbine stage 64,
66, multiple
turbine vanes 72, 74 can be provided in a ring and can extend radially
outwardly relative
to the centerline 12, while the corresponding rotating blades 68, 70 are
positioned
downstream of and adjacent to the static turbine vanes 72, 74 and can also
extend radially
outwardly relative to the centerline 12, from a blade platform to a blade tip.
It is noted that
the number of blades, vanes, and turbine stages shown in Figure 1 were
selected for
illustrative purposes only, and that other numbers are possible.
[0026] The blades 68, 70 for a stage of the turbine can be mounted to a
disk 71, which
is mounted to the corresponding one of the HP and LP spools 48,50, with each
stage having
its own disk 71, 73. The vanes 72, 74 for a stage of the compressor can be
mounted to the
core casing 46 in a circumferential arrangement.
[0027] The portions of the engine 10 mounted to and rotating with either or
both of the
spools 48, 50 are also referred to individually or collectively as a rotor 53.
The stationary
portions of the engine 10 including portions mounted to the core casing 46 are
also referred
to individually or collectively as a stator 63.
[0028] In operation, the airflow exiting the fan section 18 is split such
that a portion of
the airflow is channeled into the LP compressor 24, which then supplies
pressurized
ambient air 76 to the HP compressor 26, which further pressurizes the ambient
air. The
pressurized air 76 from the HP compressor 26 is mixed with fuel in the
combustor 30 and
ignited, thereby generating combustion gases. Some work is extracted from
these gases by
the HP turbine 34, which drives the HP compressor 26. The combustion gases are
discharged into the LP turbine 36, which extracts additional work to drive the
LP
compressor 24, and the exhaust gas is ultimately discharged from the engine 10
via the
exhaust section 38. The driving of the LP turbine 36 drives the LP spool 50 to
rotate the
fan 20 and the LP compressor 24.
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[0029] A remaining portion of the airflow 78 bypasses the LP compressor 24
and
engine core 44 and exits the engine assembly 10 through a stationary vane row,
and more
particularly an outlet guide vane assembly 80, comprising a plurality of
airfoil guide vanes
82, at the fan exhaust side 84. More specifically, a circumferential row of
radially
extending airfoil guide vanes 82 are utilized adjacent the fan section 18 to
exert some
directional control of the airflow 78.
[0030] Some of the ambient air supplied by the fan 20 can bypass the engine
core 44
and be used for cooling of portions, especially hot portions, of the engine
10, and/or used
to cool or power other aspects of the aircraft. In the context of a turbine
engine, the hot
portions of the engine are normally the combustor 30 and components downstream
of the
combustor 30, especially the turbine section 32, with the HP turbine 34 being
the hottest
portion as it is directly downstream of the combustion section 28. Other
sources of cooling
fluid can be, but is not limited to, fluid discharged from the LP compressor
24 or the HP
compressor 26. This fluid can be bleed air 77 which can include air drawn from
the LP or
HP compressors 24, 26 that bypasses the combustor 30 as cooling sources for
the turbine
section 32. This is a common engine configuration, not meant to be limiting.
[0031] Figure 2 illustrates a plurality of radially extending
circumferentially spaced
airfoils, or blades 70, with each blade 70 extending from a root 88 and
terminating in a tip
(Figure 3) arranged in a circumferential row and supported by an arcuate inner
band 96 and
an arcuate outer band 98. The arcuate outer band 98 comprises a shroud
assembly 100
made up of separate individual shroud elements 102, having opposing radial
sides 107,
109, which together circumscribe the blades 70.
[0032] Each shroud element 102 as depicted in Figures 3 and 4 is integrally
formed
with the blade 70 at the tip 90 and comprises a flange 104. The flange 104
includes a first
and a second radial edge 106, 108 and a seat 110 that projects
circumferentially beyond the
first radial edge 106 illustrated in Figure 3. The seat 110 is formed by the
first radial edge
106 and the flange 104 in a portion 112 of the flange 104 that is
circumferentially outboard
of the first radial edge 106 illustrated in Figure 4.
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[0033] Figure 5 illustrates an airfoil assembly 114 in which the shroud
element 102 is
integrally formed with the blade 70 wherein the blade terminates in a dovetail
116. The
dovetail 116 is formed to mount to the rotor 53. When assembled, the blade 70
is sprung
118 to apply a preload to the interlocks. The blade 70 can be sprung as shown
by solid line
118 from a predominantly parallel position 120 with respect to a neutral axis
122 of the
airfoil assembly 114 to a bowed position 124 when interlocked. The blade 70
can also be
sprung 118 from an initial position of predominantly bowed 126 to a parallel
position 128
when interlocked.
[0034] Regardless of the initial or final positions of the blade 70, the
final position 124,
128 will cause the second radial edge 108 bias outwardly and the seat to bias
inwardly.
This bias is caused by the compressive force Fc from the dovetail 116 which
translates to
an upward force F2 at the second radial edge 108 and a downward force Fi from
the seat
110.
[0035] Circumferentially adjacent shroud elements 102 interlock together
forming a
plurality of interlocks 130 between adjacent shroud elements 102 to form the
shroud
assembly 100 as illustrated in Figure 6. A cross-sectional view of an
exemplary
embodiment of the shroud assembly 100 of Figure 6 is shown in Figure 7 where a
first
flange 132 having a first interlock element 134 mates with a second flange 136
having a
second interlock element 138 where the first flange 132 overlies and abuts the
second
flange 136.
[0036] When the shroud assembly 100 is assembled, the second radial edge
108 of the
second flange 136 will bias toward the first flange 132 due to the forces Fi
and F2. This
bias enables friction forces to form between the first and second interlock
element surfaces
that bond each shroud element 102 to the next radially adjacent shroud element
102.
[0037] A second embodiment of the shroud assembly is contemplated in Figure
8. The
second embodiment is similar to the first embodiment, therefore, like parts
will be
identified with like numerals increasing by 100 respectively, with it being
understood that
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the description of the like parts of the first embodiment applies to the
additional
embodiments, unless otherwise noted.
[0038] In the
second embodiment a first interlock element 234 formed on a first flange
232 includes an angled seat 210 formed to receive an angled second interlock
element 238
formed on a second flange 236. The flange 204 of each shroud element 202
therefore
includes an angled seat 210 and an angled portion 240 formed to fit into the
angled seat
210. While illustrated as two ramps 242, 244 forming an apex 246, the angled
seat 210 and
angled portion 240 can be any shape where the first interlock element 234 is
formed to
receive the second interlock element 238.
[0039] A method of
forming a shroud assembly, comprising a shroud element
integrally formed with a blade, about a plurality of rotating blades in a
turbine engines
includes forming an interlock between circumferentially adjacent tips of the
blades and
preloading the interlock. The preloading of the interlock where an interlock
element is
made to bias towards another interlock element.
[0040] The
embodiments described herein have benefits regarding production,
performance, and damping capability. Prior art for shroud blade assemblies
include Z-
shaped interlocks. Implementing airfoil and shroud radial bending ensures
contact between
interlock elements to achieve outer band preload at operating conditions
without typical
torsional bending used in Z-shaped shroud design which requires a pre-twist.
This type of
bending also only requires blade balancing for centrifugal forces and improves
damping at
blade resonant vibrations due to increased contact areas between the interlock
element
surfaces. This increase in contact area also provides for a reduction in outer
flowpath
leakages. The simplified shape of design and removing a need for a pre-twist,
eases
manufacturing.
[0041] It should
be appreciated that application of the disclosed design is not limited to
turbine engines with fan and booster sections, but is applicable to turbojets
and turbo
engines as well.
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[0042] While there
have been described herein what are considered to be preferred and
exemplary embodiments of the present invention, other modifications of these
embodiments falling within the scope of the invention described herein shall
be apparent
to those skilled in the art.
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