Note: Descriptions are shown in the official language in which they were submitted.
~ 1 --
Description
Rotor With Double Pass Blade Root Cooling
Technical Field
This invention relates to gas turbine engine
5 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~be cooled during ~ 1Z~5/~
engine operation. This has typically been accom-
plished 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 extending passageways through the
enlarged rim portion of the disk. These metering
passageway6 communicate with radially extending
channels through the blade roots which feed the
hollow airfoils.
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;m; ze the amount of
cooling air ~owneeded to maintain acceptable part
operating temperatures since this improves engine
e~ficiency. It is also desirable to avoid putting
holes through the disks, since these holes weaken the
disk and limit its life.
Disclosure of Invention
An object of tha present invention is to reduce
the amount of coolant flow needed to maintain gas
turbine engine rotor blade roots and rotor blade disk
lugs within acceptable operating temperatures.
According to the present invention a turbine
rotor di~k cooperates with blade roots disposed in
slots spaced around the rim of the disk to define a
pair of cooling air passageways through each disk
slot, wherein the passageways are in series flow
relation~hip with each other such that cooling air
fl~7s in a downstream direction through one of the
passageways and thence into and through the other
passageway in the opposite direction.
In the prior art the cooliny air, after having
made one pass through the slot in the downstream
direction, still had additional cooling capacity
which went substantially unutilized. In the present
invention this relatively cool air is routed back
through the slot in an upstream direction. Twenty-six
percent less cooling air mass flow is required with
the cooling arrangement of the present invention
compared to the prior art.
In a preferred embodiment the first pass of
cooling air through the slot is through a first
passageway formed between the inner end of the blade
root and the base of the slot, wh,ich is the live rim
of the disk. Radial passageways through the blade
root, for carrying cooling air into hollow airfoils
integral with the blade root, intersect the first
passageway. A portion of the cooling air through the
first passageway is diverted into the airfoil.
The foregoing and other objects, features and
advantages of the present invention will become more
apparent in the light of the following detailed
description of preferred embodiments thereof as
shown in the accompanying drawing.
Brief Description of the Drawing
Fig. l 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
~he line 2-2 of Fig. 1.
Fig. 3 is a sectional view taken generally along
the line 3-3 of Fig. l.
Pig. 4 is a sectional view taken generall~
along the line 4-4 of Fig. 1.
Fig. 5 is a perspective view, looking generally
rear,7ard, of one segment o~ the annular rear blade
retainer for the first stage turbine rotor.
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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 ~ode for Carrying Out the Invention
As an exemplary embodiment of the present
invention consider khe ~ortion of khe turbine section
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
assembl~ 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 axiall~ through
the dis~ 16 from the disk front face 28 to the disk
rear face 30. The slots 26 are formed between what
are herein referred to as disk lugs 32. Axially
extending cooling air passagewa~s 35 are formed
between khe innermost end surface 37 of the root
end 24 and the live rim 39 of the disk 16. These
pa~3agPways 35 are for carrying cooling air khrough
th~ slot~ 26 from a fronk 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 27 have inlets 29 which
cornmunicate 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-txee 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, radiall~ 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
rim of the disk 34. A first axially extendi~g
cooling air passageway 55 is thereby formed there-
between for carxying cooling air through the disk
slot 46 from a compartment, such as the compartment
66 on th~ 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
st~ge 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 6~ is disposed radially out-
wardly of the flanged support arrns 58, 60 and
extends axially between the rear face 30 of the first
stage disk 16 and the front face 48 of the second
staye 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)
extending axially thereacross for metering a flow of
cooling air from the rear cooling air space 33 into
the intennediate compartment 66, as will be further
explained hereinbelow. Similarly, the rearward end
74 of the spacer 64 includes a radially ou~wardly
fa~ing 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. ~ plurality of
circumferentially spaced apart radial slots 75 in
the rearward end 7~ are aligned with a plurality of
circumferentially spaced apart radial slots 77 in
the front face 48 of the disk 34 to form passage-
ways for the flow of cooling air from the compart-
ment 66 into and through the rirst 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
15 is an annular blade retaining plate 90. More
specifically, the radially inner end 92 of the plate
90 includes an axially exten~; ng flange 94 having a
radially outwardly facing cylindrical surface 96.
The front ~ace 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 radially relative to the disk 16. The
plate 90 is trapped axially in position ~y a split
ring 101 and an inner annular seal carrier 102 which
is ~olted to a radially inwardly ext~n~; ng flange 104
3~
o the disk 16, such as by ~olts 105. 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,
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
~a~e 28 defining the annular cooling air space 31
~q~
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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 plurality of blade retaininy 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
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 deine the hereinabove referred to rear annular
cooling air space 33 which receives the cooling air
flowing through the ~assageways -~3 within the blade ~ ~ IZ~ 3
root slots 26. A forwardly facing, circumfer~ntially
ext~n~li ng surface 154 near the radially outermost
edge ~44 of each segment 138 bears against the disk ~ ~ IZ~/~3
face 30 (actually the lugs 32) and the end faces of
the 'ir tree shaped blade roots ~o form a full
annular seal, which seal is improved by a wire seal
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 ~nnular surace 162 of the spacer 64 and,
~3~
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along with a wire seal 164 disposed in the annular
groove defined by 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 surfacesl50 extends from
the innermost edge 144 of the segment 138 to the step
148. Slots 152, best seen in Fig. 4, ar~ thereby
formed between the abutting segments 138. The slots
152 pro-~ide fluid flow cor~munication between the gas
space 33 and the intermediate compartment 66, via
the hereinabove referred to metering cutouts 71 in
the foxward end 68 of the spacer 64. l~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~; ni ng segments 138 are supported
and positioned radially by a forwardly extending
arcuate lip 168 having a radially outwardly ~acing
sur~ace 170 which rests on a radially inwardly facing
cylindrical surface 172 o~ the disk 16. A lug 174
on each segment 138 engages a rearwaxdly 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
S 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
movemen~ by an annular rear coverplate 188. The
rear coverplate 188 has an annular, forwardly exten-
ding li~ 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
aciny ~urfaces of the disk lugs 47 and the rearward-
1~ 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 138 i9
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spaced axially from the rear face 50 of the disk 34
to ~efine the previously referred to annular gas
space 57 therebetween.
As best shown in ~igs. 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
co~erplate 178 to pro~ide fluid communication bPtween
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
cc ;cate, through the slots 75, 77, with the
intPrme~llate 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. ~ecause the pressure in the inter-
medlate compartment 66 is higher than the pressure in
the spaces 186, the cooling air ~lows from the com-
partment 66 through the first passageways 55 into the
gas space 57 and thence, in ~he opposite, orward
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direction, through the second cooling air passage~7ays
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 di~k 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 air~oil
compartrnents 215. ~he air flowing into the compart
ments 215 leaves the air~oil via holes and slots
(not 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
1O~7er 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 cooliny arrangement has been provided whereb~
cooling air from a compartmen~ upstream o~ 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 khe 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.
