Note: Descriptions are shown in the official language in which they were submitted.
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Description
Two Stage Rotor Assembly with
Improved Coolant Flow
Technical Field
This invention relates to gas turbine engine
rotors, and more particularly to rotor disk and blade
root cooling.
Background Art
In the hot, turbine section of a gas turbine
engine it is required that the roots of turbine blades
and the live rim of the turbine disk and the disk lugs
be cooled during engine operation. This has typically
been accomplished by passing cooling air across the
disk through axial passageways formed in the blade
root slot between the blade root inner end and the
disk live rim. The cooling air flow passes once through
the slot in a downstream direction and empties into
a compartment on the downstream side of the disk.
It is also usual for gas turbine engine turbine
airfoils to be "hollow"; that is, to have passageways
and/or compartments therewithin for the flow of cooling
air therethrough to maintain the airfoil temperature
below a predetermined level. It is known in the prior
art to meter a portion of cooling air from upstream
of the disk into the hollow airfoils via radially ex-
tending passageways through the enlarged rim portion
of the disk. These metering passageways communicate
with radially extending channels through the blade roots
which feed the hollow airfoils.
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In a two stage turbine, both stages are cooled
using cooling air from a compartment upstream of the
first stage disk. The cooling air for the second
stage disk rim and blades is conducted from this
upstream compartment, via axial holes in the first
disk, into an intermediate compartment formed between
the first and second stage disks. The cooling air is
then passed, for example, from the intermediate com-
partment into the hollow airfoils of the second stage
rotor via metering passageways extending substantially
radially through the enlarged rim portion of the disk.
The metering passageways communicate with channels
through the blade roots which feed the hollow airfoils.
It is desirable to m;n;~; ze the amount of
cooling air flow needed to maintain acceptable part
operating temperatures since this improves engine
efficiency. It is also desirable to avoid putting
holes through the disks, since these holes weaken the
disk and limit its life.
Disclosure of Invention
One object of the present invention is a two
stage turbine assembly with improved means for
bringing cooling air to the rims and blades of both
turbine rotors.
According to the present invention, a two stage
turbine has a first stage disk with a plurality of
circumferentially spaced apart blade root slots
extending axially therethrough about the periphery
thereof and having blades disposed therein, and a
second stage disk with a plurality of circumferentially
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spaced apart blade root slots extending axially
therethrough about the periphery thereof and having
blades disposed therein, wherein spacer means extends
between and engages the two disks defining an inter-
mediate cooling air compartment therebetween radially
inwardly of the blade root slots, said disks and
spacer means being constructed and arranged wherein
cooling air from a compartment upstream of the first
disk travels through the blade root slots of the first
disk to the rear side of the first disk and thence
radially inwardly into the said intermediate compart-
ment between the disks from whence it flows into and
through the blade root slots of the second disk to
the rear side thereof.
In a preferred embodiment, the airfoils are
hollow and a metered portion of the cooling air
passing through the blade root slots of each disk
is directed radially outwardly into internal compart-
ments of the hollow airfoils via radially extending
channels through the blade roots.
An important feature of the present invention is
the elimination of the axial holes through the first
disk which were used, in the prior art to bring
cooling air downstream to the second stage disk.
Brief Description of the Drawing
Fig. 1 is a simplified sectional view of the
turbine section of a gas turbine engine incorporating
the features of the present invention.
Fig. 2 is a sectional view taken generally along
the line 2-2 of Fig. 1.
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Fig. 3 is a sectional view taken generally along
the line 3-3 of Fig. 1.
Fig. 4 is a sectional view taken generally
along the line 4-4 of Fig. 1.
Fig. 5 is a perspective view, looking generally
rearward, of one segment of the annular rear blade
retainer for the first stage turbine rotor.
Fig. 6 is a sectional view partly broken away,
taken generally along the line 6-6 of Fig. 3.
Fig. 7 is a sectional view taken generally
along the line 7-7 of Fig. 6.
Best Mode For Carrying Out The Invention
As an exemplary embodiment of the present
invention consider the portion of the turbine section
lS of a gas turbine engine, the turbine section being
generally represented by the reference numeral 10 in
Fig. 1. Only the first two stages are shown. The
first stage rotor assembly is generally represented
by the reference numeral 12. The second stage rotor
assembly is generally represented by the reference
numeral 14.
The first rotor assembly 12 comprises a disk 16
having a plurality of blades 18 circumferentially
spaced about the periphery thereof. Each blade 18
comprises a root portion 22 and an airfoil portion
20 having a platform 25 integral therewith. With
reference also to Fig. 2, the root portion 22 has a
fir-tree shaped root end 24 disposed in a similarly
shaped fir-tree slot 26 which extends axially through
the disk 16 from the disk front face 28 to the disk
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rear face 30. The slots 26 are formed between what
are herein referred to as disk lugs 32. Axially
extending cooling air passageways 35 are formed
between the innermost end surface 37 of the root
end 24 and the live rim 39 of the disk 16. These
passageways 35 are for carrying cooling air through
the slots 26 from a front annular space 31 on the
front side of the disk 16 into a rear annular space
33 on the rear side of the disk 16 to cool the blade
root ends 24, the disk lugs 32, and the live rim
39 of the disk 16. A portion of the cooling air
flowing through the passageways 35 is diverted into
cooling air passageways or compartments 23 within the
airfoils 20 via channels 27 through the blade root
ends 24. The channels ~7 have inlets 29 which
communicate directly with the passageways 35 through
the slots 26.
The second rotor assembly 14 comprises a disk
34 having a plurality of blades 36 circumferentially
spaced about the periphery thereof. As best shown
in Figs. 1 and 3, each blade 36 comprises a root
portion 40 and an airfoil portion 38 having a plat-
form 42 integral therewith. The root portion 40
includes a fir-tree shaped root end 44 disposed in
similarly shaped fir-tree slots 46 formed between
disk lugs 47. The slots 46 extend axially through
the disk 34 from the disk front face 48 to the disk
rear face 50. The innermost, radially inwardly
facing surface 51 of each root end 44 is spaced
radially from the radially outwardly facing bottom
surface 53 of the slot 46, which is also the live
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rim of the disk 34. A first axially extending
cooling air passageway 55 is thereby formed there-
between for carrying cooling air through the disk
slot 46 from a compartment, such as the compartment
66 on the front side of the disk 34 to an annular
space 57 on the rear side of the disk 34. Further
aspects of the cooling configuration for second
stage disk and blades will be described hereinbelow.
The disks 16, 34 are connected to an engine
shaft assembly 52 through an annular support member
54 which is splined to the shaft assembly 52 as at
56. More specifically, the disk 16 includes a
flanged cylindrical support arm 58, and the disk 34
includes a flanged cylindrical support arm 60.
The flanged arms 58, 60 are secured to the support
member 54 by suitable means, such as a plurality of
nut and bolt assemblies 62.
An annular spacer 64 is disposed radially out-
wardly of the flanged support arms 58, 60 and
extends axially between the rear face 30 of the first
stage disk 16 and the front face 48 of the second
stage disk 34 defining an intermediate annular
cooling air compartment 66 radially outwardly of
the support arms and which extends axially between
the rear face 30 and the front face 48. The forward
end 68 of the spacer 64 includes a radially outwardly
facing cylindrical surface 70 which engages a
corresponding radially inwardly facing cylindrical
surface 72 of the rear face 30. The cylindrical
surface 70 includes a plurality of circumferentially
spaced apart scallops or cutouts 71 (see Fig. 4)
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extending axially thereacross for metering a flow of
cooling air from the rear cooling air space 33 into
the intermediate compartment 66, as will be further
explained hereinbelow.: Similarly, the rearward end
74 of the spacer 64 includes a radially outwardly facing
cylindrical surface 76 which engages a corresponding
radially inwardly facing cylindrical surface 78 of
the front face 48 of the disk 34. The spacer 64 is
thus supported radially by the disks 16, 34 and rotates
therewith. A plurality of circumferentially spaced
apart radial slots 75 in the rearward end 74 are aligned
with a plurality of circumferentially spaced apart
radial slots 77 in the front face 48 of the disk 34
to form passageways for the flow of cooling air from
the compartment 66 into and through the first cooling
air passageways 55 within the blade root slots 46.
In this embodiment the spacer 64 carries a
plurality of radially outwardly extending knife edges
80 which are closely spaced from a stationary annular
seal land 82. The seal land 82 is supported, through
suitable structure, from the inner ends 84 of a plurality
of circumferentially spaced stator vanes 86 disposed
between the first and second stage rotor airfoils 20,
38, respectively. The vanes 86 are supported from
an outer engine casing 88.
Secured to the front face 28 of the disk 16
is an annular blade retaining plate 90. More
specifically, the radially inner end 92 of the plate
90 includes an axially extending flange 94 having a
radially outwardly facing cylindrical surface 96.
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The front face 28 of the disk 16 includes an axially
extending flange 98 having a radially inwardly
facing cYlindrical surface 100. The surface 96
mates with the surface 100 to orient and support the
plate 90 radiallv relative to the disk 16. The
plate 90 is trapped axially in position by a split
ring 101 and an inner annular seal carrier 102 which
is bolted to a radially inwardly extending flange 104
of the disk 16, such as by bolts 106. The seal
carrier 102 includes a plurality of conventional,
radially outwardly extending knife edges 108 which
are in sealing relationship to a stationary annular
seal land 110 secured to stationary structure gen-
erally represented by the reference numeral 112.
The plate 90 also include an axially extending
cylindrical seal carrier 114 integral therewith and
which carries a plurality of conventional, radially
outwardly extending knife edges 116. The knife
edges 116 are in sealing relationship with a
stationary annular seal land 118 secured to the
stationary structure 112. The stationary structure
112 cooperates with a stage of stator vanes 120
disposed in the gas path upstream of the rotor
blades 20. The vanes 120 are secured by suitable
means to the engine outer case 88.
The plate 90 further includes a frusto-conical
portion 126 extending radially outwardly in a down-
stream direction. The frusto-conical portion 126
has a radially outer end 128. The end 128 includes
an annular surface 61 facing axially downstream
which abuts the front face 28 of the disk 16 and the
fir-tree shaped blade root ends 24. With reference
to Fig. 1, the seal carriers 102, 114, the plate 90,
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and the stationary structure 112 define an inner
annular compartment 122 which is fed cooling air
from a plurality of circumferentially spaced apart
nozzles 124. The plate 90, between its inner and
outer ends 92, 128, stands away from the disks front
face 28 defining the annular coolinq air space 31
which, through large holes 132 in the plate 90, is
in fluid communication with and is, in effect, a part
of the compartment 122. The knife edges 116 and a
wire seal 134 between the plate end 128 and disk face
28 prevent leakage from the compartments 122, 31
radially outwardly into an outer gas space 136.
Secured to the rear face 30 of the first disk
16 are a plurali.y of blade retaining segments 138
circumferentially disposed about the engine axis.
One of these blade retaining segments 138 is shown
in perspective in Fig. 5. Each segment 138 includes
oppositely facing end surfaces 140, 142. The end
surfaces 140 abut the end surfaces 142 of adjacent
2~ segments to form a segmented full annular member.
The segments 138 are trapped axially between the
spacer 64 and the rear face 30 of the first disk 16
to define the hereinabove referred to rear annular
cooling air space 33 which receives the cooling air
flowing through the ~assageways ~ within the blade ~ ~z-~S-83
root slots 26. A forwardly facing, circumferentially
extending surface 154 near the radially outermost
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~edge ~t* of each segment 138 bears against the disk
~'l face 30 (actually the lugs 32) and the end faces of
the fir tree shaped blade roots to form a full
annular seal, which seal is improved by a wire seal
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156 disposed in an annular groove formed by arcuate
groove segments 158 in each of the blade retaining
segments 138. Similarly, rearwardly facing arcuate
surface segments 160 bear against the forwardly
facing annular surface 162 of the spacer 64 and,
along with a wire seal 164 disposed in the annular
groove defined ~y arcuate groove segments 166
(Fig. 5), form a full annular seal against the
surface 162.
Each end face 140, 142 is cut back or stepped,
as at 148, such that a surface 150 is formed parallel
to but is out of the plane of its respective end
surface 140, 142. The surfaces150 extends from
the innermost edge 144 of the segment 138 to the step
148. Slots 152, best seen in Fig. 4, are thereby
formed between the abutting segments 138. The slots
152 provide fluid flow communication between the gas
space 33 and the intermediate compartment 66, via
the hereinabove referred to metering cutouts 71 in
the forward end 68 of the spacer 64. ~etering
holes 151 (Fig. 4) formed between abutting segments
138 provide fluid flow communication between the gas
space 33 and outer annular compartment 153. The
cooling air flowing into the compartment 153 is
used to cool the knife edges 80 and seal land 82.
The blade ret~; n; ng segments 138 are supported
and positioned radially by a forwardly extending
arcuate lip 168 having a radially outwardly facing
surface 170 which rests on a radially inwardly facing
cylindrical surface 172 of the disk 16. A lug 174
on each segment 138 engages a rearwardly extending
annular flange 176 of the disk 16 to further position
the segments 138 both axially and radially relative
to the disk 16.
The second stage disk 34 also includes blade
retaining means on both the front and rear sides
thereof. In this embodiment, the spacer 64 is also
the front side blade retainer. More specifically,
the rearward end of the spacer 64 includes a
radially outwardly extending annular coverplate 178
having a rear surface 180 which abuts the front
surfaces of the lugs 47 and the front surfaces 182
of the blade root portions 40. These front surfaces
are substantially coplanar. The coverplate 178
extends radially outwardly to the blade platforms 42
such that it completely covers or closes off the
forward end of the space or volume 186 defined
between the extended portions 187 of the root portions
40.
The blades are prevented from axially rearward
movement by an annular rear coverplate 188. The
rear coverplate 188 has an annular, forwardly exten-
- ding lip 190 which snaps over a shoulder 192 on the
rear side of the disk 34 thereby supporting and
positioning the coverplate radially. The rear
coverplate is trapped axially by a split annular
ring 193 which engages the radially innermost end
of the coverplate 188 and fits tightly between it and
a radially outwardly extending annular flange 194
of the disk 34. The radially outermost end 196 of
the coverplate 188 includes a forwardly facing
annular surface 198 which forms an annular seal
against the substantially coplanar rearwardly
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facing surfaces of the disk lugs 47 and the rearward-
ly facing surfaces of the blade root ends 44.
Between the snap diameter at the shoulder 192 and the
seal at the surface 198 the cover plate 188 is
spaced axially from the rear face 50 of the disk 34
to define the previously referred to annular gas
space 57 therebetween.
As best shown in Figs. 3 and 6, the radially
inwardly facing surfaces 200 of the outer teeth 202
of the root portion 40 are spaced radially outwardly
from the corresponding opposed surfaces 204 of the
disk lug inner teeth 206 to define second air
cooling passageways 208 through the slots 46. These
passageways have inlets 209 at the rear face 50 of
the disk 34 which communicate with the gas space 57.
The radially outermost portion of the front face
of each lug 47 is cut back slightly as at 210 so as
to be spaced slightly from the surface 180 of the
coverplate 178 to provide fluid communication between
outlets 211 of the second cooling air passageways
208 and the spaces 186 between the blade root
portions 40.
The first cooling air passageways 55 have
inlets 212 and outlets 214. The inlets 212
communicate, through the slots 75, 77, with the
intermediate cooling air compartment 66 between the
first and second rotor disks 16, 34. The outlets
214 open into the gas space 57 on the rear side of
the disk 34. The first and second passageways 55,
208 are in series fluid flow relation through the
gas space 57. Because the pressure in the inter-
mediate compartment 66 is higher than the pressure in
the spaces 186, the cooling air flows from the com-
partment 66 through the first passageways 55 into the
gas space 57 and thence, in the opposite, forward
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direction, through the second cooling air passageways
208. The air then flows into the spaces 186 via the
cutouts 210 in the lugs 47. From the spaces 186 the
cooling air travels into another compartment (not
shown) located downstream thereof. The cutouts 210
are sized to meter the flow of cooling air through
the blade root slots 46.
Referring to Figs. 6 and 7, in a preferred
embodiment, the second stage airfoils 38 have cooling
air passageways or compartments 215 therein which are
fed cooling air from the intermediate compartment 66
between the disk 16, 34 via a radially extending
channel 216 through the blade root portion 40.
The channel 216 interconnects the airfoil compartments
215 and the first cooling air passageway 55 through
the root slot 46. An inlet 218 to the channel 216 is
covered by a thin plate 220. The plate 220 has a
metering orifice 222 therethrough aligned with the
channel inlet 218 for metering the appropriate amount
of flow from the first passageway 55 into the airfoil
compartments 215. The air flowing into the compart-
ments 215 leaves the airfoil via holes and slots
tnot shown) through the airfoil wall for cooling the
same, as is well known in the art. During rotor
operation, the pressure in the compartments 215 is
lower than the pressure in the intermediate cooling
air compartment 66 such that the airflow is in the
proper direction.
Considering the turbine section 10 as a whole,
a novel cooling arrangement has been provided whereby
cooling air from a compartment upstream of the first
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stage rotor disk 16 is used to cool the first and
second stage disk lugs, live rims, blade roots and
airfoils. This turbine section construction is
particularly unique in that it requires no life
limiting holes through the first stage disk to get
cooling air from upstream thereof to the second stage
blade roots and into the second stage airfoils 38.
Furthermore, the unique double pass cooling air flow
arrangement through the second stage blade root area
reduces the cooling air mass flow requirements for
cooling the second stage disk rim, lugs and blade
roots by twenty-six percent (26%).
Although the invention has been shown and
described with respect to a preferred embodiment
thereof, it should be understood by those skilled in
the art that other various changes and omissions in
the form and detail thereof may be made therein
without departing from the spirit and the scope of
the invention.
