Note: Descriptions are shown in the official language in which they were submitted.
13DV-8814
CONTOURED ROTOR BLADE SHROUD SEGMENT
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Field of the Invention
This invention relates to gas turbine
engines, and, more particularly, to improved shroud
segments mounted to the casing of the high pressure
turbine or compressor of a gas turbine engine such as
a jet engine.
Backqround of the Invention
The turbines and compressors of gas turbine
engines such as jet engines each include one or more
circumferentially extending rows or stages of rotating
rotor blades which are axially spaced between rows or
stages of fixed stator vanes. Each rotor blade has a
blade root mounted to the rotor disk, and an air foil
extending radially outwardly from the root which
terminates at a blade tip. In many gas turbine engine
designs, a number of abutting, circumferentially
extending shroud segments are carried by the turbine
or compressor case to form an essentially continuous
cylindrical-shaped surface along which the tips of the
rotor blades tangentially pass. Each of these shroud
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segments includes an outer face, and an inner,
arcuate-shaped face along which the blade tips pass,
opposite end portions which abut with adjacent shrouds
and opposed side mounting rails which mount to sta-
tionary hangers on the casing of the turbine and/orcompressors.
The shroud segments, particularly those
located in the turbine of a jet engine, are subjected
to high temperatures at their inner face along which
the rotor blades pass. In an effort to lower the
temperature of the shroud segments and increase their
durability, cooling air from an intermediate stage of
the compressor is often directed onto the outer face
of the shrouds. This cooling air is intended to
raduce the overall temperature of the entire shroud
without directly contacting the inner face and dis-
rupting the air flow through the turbine or compres-
sor.
A major design consideration in any jet
engine is the reduction of specific fuel consumption.
one source of decreased specific fuel consumption in
many jet engine designs is pressure losses resulting
from the creation of a relatively large radial tip
clearance between the tip of the rotor blades and the
inner face of the shroud segments. It is believed
that one source of increased radial tip clearance is
attributable to a problem known as "chording".
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Chording results from the temperature differential
between the high temperature inner face and the cooler
outer face of the shroud segments. Impingement of
cooling air on the outer face of the shroud segments
while the inner face is subjected to high temperatures
causes the shrouds to chord or "straighten out"
circumferentially, i.e., the end portions of the inner
face of the shroud tend to move radially outwardly
relative to the center portion of the inner face of
the shroud. While the interconnection of the side
mounting rails of the shroud segments with the sta-
tionary hooks on the case of the compressor or turbine
is intended to resist chording or straightening-out of
the shroud segments, such resistance is overcome by
the temperature gradient between the outer and inner
faces thereof.
Because the inner face of these shroud
segments is formed with a substantially constant
radius of curvature from end~to-end, chording has the
effect of creating a wedge-shaped space or gap between
the tip of the rotor blades and each end portion of
the inner face of the shroud segments. Such chording
can also cause additional blade tip rubs in the
central portion of the shroud segment inner surface.
These rubs produce friction which further increases
the radial temperature gradient, thereby causing even
further chording and rubs. This increase in radial
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tip clearance at the end portions of each shroud
segment has been found to be equivalent to a uniform
tip clearance increase of about 0.00~ inches in some
types of jet engines, resulting in pressure losses
which reduce specific fuel consumption by a signifi-
cant amount, e.g., about 0.4%.
One attempt to reduce chording, or the
straightening out of the shroud segments, has been to
form one or more radially extending notches or grooves
in each of the side mounting rails of the shroud
segments which mount to stationary structure of the
turbine or compressor casing. These radial grooves
are intended to reduce or eliminate the "beam
strength" of the shroud segments by making them
lS discontinuous along the length of their side mounting
rails.
It has been found that the presence of
radial notches or grooves in the shroud segments
creates high stress concentrations at the inner end of
such grooves. These stress concentrations can create
cracking or fracturing of the shroud segments which
can propagate from the groove and result in premature
failure of the shroud segment.
SummarY of the Invention
It is therefore among the objectives of this
invention to provide shroud segments adapted to mount
to the casing of the turbine or compressor of a gas
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turbine engine, such as a jet engine, which impro~e
specific fuel consumption of the jet engine and which
exhibit improved durability.
These objecti~es are accomplished in a
shroud segment having a shroud body including opposed
ends, an outer face, an inner face and forward and aft
side mounting rails which are adapted to mount to
fix~d support structure on the casing of a turbine or
compressor of a jet engine. The inner face of each
shroud segment has a center portion located between a
pair of end portions, each of which extend from the
opposite ends of the shroud body toward the center
portion of the inner face. The inner face is formed
in a generally concavely arcuate shape in which each
lS f the end portions are formed with a first radius of
curvature, and the center portion is formed with a
second, smaller radius of curvature. This configura-
tion of the inner face of the shroud body is intended
to take into account "chording" of the shroud segment
during operation of the gas turbine engine so as to
avoid the formation of a large radial ~ip clearance
between the rotor blade tips and the shroud segments.
This invention is predicated upon the
concept of forming a shroud segment with an inner face
which, during operation of the turbine or compressor
of a gas turbine engine, undergoes chording and is
deflected into a position which closely conforms to
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the path of motion of the blade tips of the rotor
blades of such engine. As mentioned above, "chording"
refers to beam-bending of the shroud segment resulting
from the creation of a temperature gradient or differ-
ential between the inner and outer surfaces of theshroud segment. Cooling air is directed onto the
outer face of the shroud segment while its inner face
is subjected to the relatively high temperatures of
the turbine or compressor of the gas turbine engine.
This temperature differential creates bending forces
which are transmitted along the length of the shroud
segment and which force the outer ends of the inner
face to move radially outwardly with respect to the
center of the inner face. In other shroud segment
desiyns, this radially outward movement of the end
portions of the inner face of the shroud segment
creates a wedge-shaped gap or clearance between the
rotor blade tips and shroud segments during operation
of the gas turbine engine.
Using numerically controlled grinding
equipment, the inner face of the shroud segment of
this invention is formed with two different radii of
curvature to account for chording of the shroud
segment. The end portions of the inner face are
ground with a first radius of curvature while the
center portion of the inner face is ground with a
second, smaller radius of curvature. ~hen the shroud
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segments become hot during operation of the gas
turbine engine, "chording" or bending of the outer
portions of the inner face with respect to its center
portion results in the formation of an inner face
having a combined or deflected radius of curvature
which is approximately equal to the path of travel of
the rotor blade tips of the gas turbine engine. In
other words, as the end portions of the inner face of
the shroud segment are deflected radially outwardly
with respect to the center portion ther~of, a new
radius of curvature is formed with the end portions in
this deflected position which closely conforms to the
path of travel of the rotor blade tips.
The configuration oE the inner face of the
shroud segments of this invention therefore provides a
substantially constant and relatively small radial tip
clearance during operation of the gas turbine enyine.
This reduces pressure losses in the turbine or com-
pressor and thus improves the specific fuel consump-
tion of the gas turbine engine. No attempt is made toresist chording of the shroud segments in this inven-
tian, and therefore radial slots used in some other
shroud segment designs to reduce chording can be
eliminated which reduces or eliminates the formation
of stress concentrations in the shroud segments ~hich
had been a problem in other designs.
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Descri~tion of the Drawings
The structure, operation and advantages of
the presently preferred embodiment of this invention
will become further apparent upon consideration of the
following description, taken in conjunction with the
accompanying drawings, wherein:
Fig. 1 is a schematic cross sectional view
of a portion of the turbine of a jet engine illustrat-
ing the mounting of the shroud segment to the turbine
case;
Fig. 2 is a perspective view of a shroud
segment of this invention;
Fig. 3 ls a side cross sectional view of a
group of three shroud segments in an undeflected
position with a pair of turbine blades moving there-
pas~; and
Fig. 4 is a view similar to Fig. 3 except
with the shroud segments in a deflected position.
Detailed Descript~on of the Invention
Referring now to Fig. 1, a shroud segment
10 in accordance with this invention is shown in
position within the turbine 12 of a gas turbine
engine such as shown, for example, in United States
Patent No. 4,177,004 issued December 4, 1979,
Riedmiller et al. The turbine 12 is shown for
purposes of illustrating the positioning of shroud
segment 10, and it should be understood that the
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shroud segment 10 could be utilized in turbines of
other designs and/or within the high pressure com-
pressor of a gas turbine engine. The detailed con-
struction of the turbine 12 forms no part of this
invention per se and is thus described briefly herein.
A first stage stator vane 14 is bolted at
its inner band 16 to a first stage support 18 which
provides both radial and axial support for the stator
vane 14. An outer band 20 carried by the outside
diameter of the stator vane 14 is mounted by a ring 22
to a vane support 24. As used herein, the term
"outer" refers to a direction toward the top of Fig.
1, and the term "inner" refers to the opposite direc-
tion.
The stator vane 14 is cooled by compressor
discharge air which enters a forward plenum 26 Aefined
on its outer side by a compressor rear frame 28 and on
its inner side by an impingement plate 30. The term
"forward" as used herein refers to the lefthand side
of Yig. 1, and "aft" refers to the righthand side of
Fig. 1. The impingement plate 30 is secured by a
plurality of bolts 32 to a seal 34, and this seal 34
is secured to the vane support 24 by fastener 36. The
seal 34 is annular in shape and extends radially
outwardly from the vane support 24 to a pad 38 on the
compressor rear frame 2~ to isolate the forward plenum
26 from an aft plenum 40 which contains cooling air at
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a lower pressure and temperature from that of forward
plenum 26.
A first stage of rotor blades 42, tangen-
tially rotatable on a rotor disk 44 are located aft of
the stator vanes 14 and each have a blade tip 46
immediately adjacent the shroud segments 10. The
shroud segments 10 are supported on the stationary
structure of the turbine 12 such that a relatively
small radial tip clearance 48 is maintained between
the inner face of the shroud segments 10 and the blade
tip 46 of the rotor blades 42, as described below.
Support for the shroud segments 10 is
provided on the forward end by a plurality of shroud
support plates 50 which are connected to the vane
support 24 by the fastener 36. Each shroud support
plate 50 is formed with an a-xial flange 52 for mount-
ing the forward side of the shroud segment 10, as
described below. Structure for supporting the aft
side of shroud segment 10 includes a rim 54 integrally
formed with the vane support 24 having an outer flange
56 and an inner flange 58. This rim 54 mounts a
C-clamp 60 havinq an outer flange 62 which engages the
outer flange 56 of rim 54, and an inner flange 64
which mounts the shroud segment 10 as described below.
The shroud segment 10 is cooled by cooling
air discharged from the compressor (not shown) of the
gas turbine engine which enters a cavity 66 formed by
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the vane support 24, rim 54 and shroud support plates
50. The cooling air enters the cavity 66 through
apertures 31 formed in the impingement plate 30,
through apertures 23 of ring 22 and an opening 68 in
the shroud support plates 50. The cooling air
impinges upon the outer surface or face of the shroud
segment 10 to reduce the overall temperature of the
shroud segment 10.
Referring now to Fig. 2, one of the shroud
segments 10 is illustrated in detail. Each shroud
segment 10 includes a shroud body formed with a
forward side mounting rail 72, an aft side mounting
rail 74, a center stiffener 75 extending between the
side rails 72, 74, opposed end plates 76, 78, an outer
face 80 and an inner Pace 82 described in detail
below.
The forward and aft side mounting rails 72,
74 each comprise an outer arm 84 and an inner arm 86
which are ~oth connected at one edge to a side plate
88. The arms 84, 86 of the forward side mounting rail
72 are spaced from one another to form a forward slot
90 which is adapted to receive the axial flange 52 of
shroud support segment 50. Similarly, the arms 84, 86
of the aft side mounting rail 74 are formed with an
aft slot 92 therebetween which is adapted to receive
the inner flange 6~ of the C-clamp 60. The shroud
se~ment 10 is thus mounted to the stationary structure
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of the turbine 12 and a plurality of circumferentially
extending shroud segments 10 a~ut one another at their
end plates 76, 78 to form a substantially continuous
cylindrical-shaped surface consisting of adjacent
S inner ~aces 82 of abutting shroud segments lo.
As best shown in Figs. 3 and 4, the inner
face 82 of each shroud segment 10 is formed with a
center portion 94 and two end portions 96, 98 which
extend toward the center portion 94 from the end
plates 76, 78, respectively, of the shroud body. As
shown in Fig. 3, wherein the inner face 82 is in an
undeflected position, the end portions 96, 98 of the
inner face 82 are formed with a first radius of
curvature which is substantially coincident with the
path 100 of movement of the blade tips 46 of rotor
blades 42. The center portion 94 of the inner face 82
of shroud segment 10 is formed with a second radius of
curvature, which is smaller than the first radius of
curvature of end portions 96, 98, so that a gap or
space 102 of about .005 to .olo inches is formed
between the center portion 94 and the rotor blade path
100. It is contemplated that these grinding radiuses
could be formed in the shroud segment 10 by commer-
cially available computer numerically controlled
2S grinding equipment or any other machining technique.
The purpose o~ forming the inner face 82 o f
the shroud segment 10 with two radii of curvature is
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to account for "chording", i.e., the radially outward
deflection of the end portions 96, 98 of inner face
82, and the radially inward deflection of the center
portion 94 of inner face 82, which results from the
temperature differential between the cooled outer face
80 and the hot inner face 82 of shroud segment 10. As
discussed above, the inner face 82 undergoes deflec-
tion in response to this temperature differential
because the shroud segment 10 acts as a beam and
transmits bending forces along the length thereof.
As illustrated in Fig. 4, this bending or
chording of the shroud segment 10 is accommodated by
the configuration of the inner face 82. With the
shroud segment 10 in a heated, operating condition,
the end portions 96, 98 of the inner face 82 deflect
radially outwardly and the center portion 94 thereof
deflects radially inwardly such that the combined,
deflected radius of curvature of the inner face 82 of
shroud segment 10 is substantially equal to the path
~o 100 of motion of the rotor blade kips 46. In this
deflected position of the inner face 82, a substan-
tially uniform and relatively small gap 104 is created
between the rotor blade tips 46 and inner face 82
which substantially reduces pressure losses in the
turbine or compressor of the gas turbine engine.
While the invention has been described with
reference to a preferred embodiment, it will be
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understood by those skilled in the art that various
changes may be made and equivalents may be substituted
for elements thereof without departing from the scope
of the invention. In addition, many modifications may
be made to adapt a particular situation or material to
the teachings of the invention without departing from
the essential scope thereof. Therefore, it is
intended that the invention not be limited to the
particular embodiment disclosed as the best mode
contemplated for carrying out this invention, but that
the invention will include all embodiments falling
within the scope of the appended claims.
