Note: Descriptions are shown in the official language in which they were submitted.
11351~ t
PARTIALLY SEGMENTED SUPPORTING AND SEALING
STRUCTURE FOR A GUID~ VANE ARRAY OF A GAS
TURBINE ENG_NE
Background of the Invention
The present invention relates to a turbine nozzle as
employed in a multi-stage turbine of a gas turbine engine,
and more partlcularly, a supporting and sealing structure
for an array of radially extending guide vanes of a tur-
bine nozzle wherein the root ends of the guide vanes are
flexibly connected by a segmented inner shroud, and which
supporting structure includes a pressure dam to minimize
the amount of leakage introduced by providing flexibility
in the nozzle assembly.
In a multi-stage turbine of a gas turbine engine,
stationary Vane assemblies are inserted between the rotor
wheels, as well as at the entrance and exit of the turbine
unit. In the operation of the gas turbine engine, the sta-
tionary vane assemblies function to alter the static pres-
sure and change the velocity of the high pressure, high
temperature gases flowing through the turbine. Heretofore,
in order to insure the structural integrity of a vane as-
sembly ~s it is subjected to thermal excursions of the com-
ponents of the assembly during transient and steady state
operating conditions of the gas turbine engine, it has been
common to cast the entire nozzle ass~mbly in one piece. The
one piece assembly included an outer, unitary shroud, an
inner, unitary shroud, and the array of radially extending
guide vanes. With this prior art construction, it has been
found that during transient and steady state operation of
the gas turbine engine, the temperature differentials be-
tween the thin, fast responding vanes and the slower, more
massive shroud rings, causes a differential thermal growth
1~3519S
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or thermal gradient to develop within the nozzle assembly,
as well as different temperature levels throughout the noz-
zle assembly. The result of the differential thermal grad-
ients causes differential thermal excursions of the parts
of the nozzle assembly, thereby leading to the development
of local stresses and cracks in the interconnections be-
tween the vanes and the shrouds. In addition, the inner
shroud of a stationary turbine nozzle is usually sealed by
; a sheet metal member which is usually brazed to the inner
shroud, and it has been found that the thermal excursions
of the parts of the turbine nozzle have caused distortion
and separation of the brazed connections due to the ther-
mal loading on the sheet metal pieces, thereby resulting in
pressure leakage through the Vane assembly.
Accordingly, it is an object of the subject invention
to overcome the shortcomings of the prior art turbine nozzle
assemblies and to provide a new and improved supporting
and sealing structure for an array of radially extending
guide vanes of a nozzle of a gas turbine engine, which
supporting and sealing structure provides a flexible coup-
ling between the individual vanes and the inner shroud.
It is another object of the present invention to pro-
vide a new and improved supportlng and sealing structure
for an array of radially extending guide vanes of a gas
turbine engine wherein the flexible coupling between the
root ends of the vanes and the inner segmented shroud is
sealed by a flexible, pressure dam to minimize leakage
through the flexible coupling.
It is a further object of the present invention to
provide a new and improved supporting and sealing structure
for an array of radially extending guide vanes of a nozzle
of a gas turbine engine including means for maintaining the
radial and axial alignment of the vanes under transient and
steady state operating conditions of the gas turbine engine.
Summary of the Invention
The nozzle of the subject invention is embodied in a
gas turbine engine, and includes a radially inner shroud
ring, a radially outer shroud ring, and a plurality of rad-
ially extending Vane structures respectively disposed be-
1135195
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t~een the xadially lnner and the radially outer shroud
rings. Each vane is firmly secured at its tip end to the
radially outer shroud ring, while the root end of each vane
is secured to the inner shroud ring by an inner support and
sealing structure. The latter includes a radially inner
ring structure of generally L-shaped cross-section includ-
ing a generally cylindrical ~ase, and a radially outwardly
extending disc. The root end of each vane is connected to
a structural segment which forms a portion of a plurality
of segments defining a segmented ring. Each structural
segment includes a radially inwardly extending lug which is
adapted to engage a cooperating slot which extends in two
mutually perpendicular directions on the radially outwardly
extending disc of the inner ring structure to define a slip
fit connection. The latter functions to retain the inner
support ~nd sealing structure concentric to the outer
shroud ring. The sllp fit connection between the struc-
tural segments and the disc also functions to define a
pressure dam for minimizing pressure leakage through the
flexible coupling of the vanes to the inner shroud. A
spring of generally C-shaped cross-section preferably ex-
tends between each segment and thè base of the L-shaped in-
ner ring structure, thereby providing a flexible restrain-
; ing interconnection between the inner shroud and the vanes.
The new and improved sealin~ and supporting structure ofthe subject invention provides flexibility in the nozzle
assembly, thereby eliminating the development of local
stresses within the nozzle assembly, while minimizing the
amount of leakage introduced by providing flexibility in
the nozzle assembly. The flexibility of the sub~ect in-
vention is obtained by the provision of the segmented in-
ner shroud and the springs. The pressure dam is effective
to reduce leakage, and by virtue of the slip fit intercon-
nection between the structural segments and the disc por-
tion of the inner ring structure, the pressure dam is main-
tained during thermal excursions of the components of the
nozzle assembly, during both transient and steady state
operating conditions of the turbine engine.
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Description of the Drawings
Other objects and advantages of the invention will be-
come apparent from a reading of the following detailed de-
scription taken in conjunction wlth the drawings in which:
FIG. 1 is a front elevational view of the new and im-
proved nozzle assembly of the subject invention;
FIG. 2 is a cross-sectional view taken along line 2-2
in FIG. l;
FIG. 3 is a rear elevational view of the new and im-
proved nozzle assembly of the subject invention; and
FIG. 4 is a cross-sectional view taken along line 4-4
in FIG. 2.
Detailed Descri tion of the Preferred Embodiment
p
Referring to FIGS. 1, 2, and 3, the stationary turbine
nozzle assembly of the subject invention is generally des-
ignated by the numeral 10 and basically comprises a radi-
ally outer shroud ring 12, a radially inner shroud ring 14,
and an array of radially extending guide vanes 16 disposed
between rings 12 and 14. The radially outer tip portions
18 of each guide vane 16 is secured to the inner surface of
the outer shroud ring 12 by a rigid connection, such as by
brazing or castin~. On the other hand, the root portion 20
of each guide vane 16 is flexibly connected to the inner
shroud by means of the supporting and sealing str~cture of
the subject invention. The supporting and sealing struc-
ture enables the guide vanes 16 to undergo thermal excur-
sions during transient and steady state operation of the
gas turbine engine, without resulting in distortion or the
development of local stresses on the assembly 10 which
could lead to the development of local cracks in the assem-
bly.
The supporting and sealing structure of the subject
invention includes a segmented ring 30 which is defined by
a plurality of individual segments 32 arranged concentri-
cally with the radially outer shroud ring 12. Each segment32 is connected to the root end 20 of a radial~y extending
guide Vane 16. As illustrated in FIGS. 2, 3, and 4, depend-
ing from each segment 32 and extending radially inward of
the segment 32, is a T-shaped lug portion 34. Each T-
1135195
--5--shaped lug 34 includes a leg portion 36 which is aligned
with the longitudinal axis of the gas turbine engine, and
a transverse bar segment 38 extending orthagonal to the
longitudinal axis of the engine. The supporting and seal-
ing structure 40 further includes a radially inner ringsupport structure 40 which is generally L-shaped in cross-
section (see FIG. 2) and includes a gènerally cylindrical
base 42 and a radially outward extending disc portion 44.
Secured to the disc portion 44 is an angled ring member 50
which includes an array of radially extending cut-outs 52
so as to define a generally scalloped configuration, as
viewed from the rear of the assembly 10 (see FIG. 3). The
angled cross-section of the ring 50 (see FIGS. 2 and 4) re-
sults in a circumferential space or slot 60 extending about
lS the radially outer diameter of the disc portion 44 of the
ring support structure 40. As illustrated, the circumfer-
ential slot 60 is downstream of the disc portion 44.
The leg portions 36 of the T-shaped lugs 34 are re-
spectively slidably mounted in the cut-outs 52, while the
transverse bar segment 38 of each lug 34 is slidably mount-
ed in the space 60 defined between the disc 44 and the
angled ring 50 (see FIGS. 2 and 4). By this arrangement,
a slip fit interconnection is defined between each segment
32 and the inner ring support structure 40, with the slip
fit connection effectively maintaining the continuity be-
tween the segmented ring 30 and the inner support ring 40
so as to define a pressure dam for minimizing pressure leak-
age through the flexible coupling of the supporting and
sealing structure. It is noted that the pressure dam is
maintained throughout the various transient and steady
state operating conditions of the gas turbine engine, dur-
ing which time the thermal excursions of the vanes cause
the segments 32 to move relative to the inner support ring
40.
Disposed at the upstream end of each segment 32 and ex-
tending between said segment 32 and the upstream end of the
base 42 is a spring means in the form of a C-shaped, flex-
ible spring 70. As shown in FIG. 1, a plurality of springs
70 are provided preferably corresponding to the number of
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segments 32 of the segmented r~ng 30. Each spring 70 is
connected at its opposite ends to a segment 32 and to the
base 42 of the inner support ring 40. By this arrangement,
the springs 70 provide a constant biasing force for main-
taining the guide vanes 16 in axial and radial alignmentduring both transient and steady state operating conditions
of the gas turbine engine when the stationary vane assembly
10 and the components thereof are subjected to thermal ex-
cursions. Accordingly, the arrangement of springs 70, seg-
mented ring 30, and the inner ring support 40 effectivelydefines a flexible coupling as part of the supporting and
sealing structure of the subject invention. Furthermore,
axial positioning of the guide vanes 16 is assured by vir-
tue of the slip fit interconnection between the T-shaped
lugs 34 and the inner ring support structure 40, and in
particular, the interconnection between the transverse bar
segments 38 of the lugs 34 and the circumferential slot 60
defined between the disc 44 and angled ring 50.
In operation, the supporting and sealing structure 30
insures that the required sealing of the pressure upstream
of the vane assembly is maintained relative to the differ-
ential pressure downstream of the vane assembly, and by
virtue of the flexible coupling interconnection, differen-
tial thermal expansion and excursions of the shrouds and
the guide vanes is readily accommodated without the devel-
opment of local stresses which could lead to cracks in the
assembly 10.
Accordingly, the subject invention provides a par-
tially segmented turbine nozzle having a flexible support
and sealing inner shroud mem~er which is effective to ac-
commodate and neutralize thermal excursions of components
of the nozzle during transient and steady state operating
conditions of the gas turbine engine. The flexible coup-
ling at the inner shroud of the subject nozzle assembly in-
sures that the structural integrity of the fixed, usuallybrazed, connections of the vane tips to the outer shroud is
maintained. Furthermore, the subject construction elimin-
ates local stress problems brought about by differential
thermal expansions of the components of the assembly.
113Sl~
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Flexibility of the subject turbine nozzle is achieved by
the arrangement of segmenting the inner shroud and the pro-
vision of the springs which maintain the radial positions
of the inner ring structure 40, while providing flexibility
of the guide vanes in the radial direction. The pressure
dam forming a portion of the supporting and sealing inner
shroud construction is effective to reduce leakage through
the segmented inner shroud, and the pressure dam includes
the slip fit construction so as to maintain the pressure
dam during various operating conditions of the gas turbine
engine, while enabling free movement of the guide vanes in
the radial direction. Still further, the specific con-
struction of the pressure dam of the subject invention func-
tions to maintain and locate the axial position of the seg-
mented ring, and the slip fit construction further aids inmaintaining concentricity of the inner shroud.
Although the invention has been described with respect
to a preferred embodiment, it is readily apparent that
those skilled in the art will be able to make numerous mod-
ifications of the exemplary embodiments without departingfrom the spirit and scope of the invention. All such modi-
fications are intended to be included within the spirit and
scope of the invention as def~ned by the appended claims.
