Note: Descriptions are shown in the official language in which they were submitted.
RAc~GRotJMn or~ T~ INV~h~TION
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Field of the Invention:
This invent~on relates t,o a gas turbine blade and
more specifically to an inte~rally cast turbine blade having
lnternal flow paths for coolant flow therethrough.
Description of the_Prior Art:
In increasing gas turbine efficiencies by in-
creasing the inlet temperatures, gas turbine designers are
limited by the strength o~ materials at these ele$ated
temperatures. This is particularly critical in the rotor
blades which, in addition to being directly in the flow path
of the motive gas, are sub~ected to centrifugal and bending
forces. Thus~ ln order for the blade, even when composed o~
high strength metal alloys, to have a sufficiently long
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operating life it ls necessary to cool them to a temperature
well below that of the motl~e gas. Ilowever, ln that the
cooling fluid itself introduces ineffi.ciencies into the gas
turbine cycle, lt is also important that the coolant i9
utilized to a maximum extent before being exhausted lnto the
motive fluid rlOw patn.
U.S. Patent No. 3,533,711 shows a gas turblne
blade having two serpentine flow paths for directing a
coolant therethrough. ~owever, the leading edge of this ~
10 blade, in addition to being cooled via impingement of the ~ -
coolant thereon, contains openings through which the coolant ~`~
flows. These openings, although beneficial ln cooling the
blade, tend to cause stress concentration in a crltical area ~r -
of the blade, namely the leading edge. The present lnventlon `
provides a somewhat slmilar blade, however ef~icient leading `-
edge cooling is provided without introducing stress concen~
trating openings, with the coolant being dlrected in such a
manner that that area of the leading edge which is generally
weakest, because its exposure to one of the higher tempera~
ture areas of the hot motive gases in combination with the
centrlfugal force induced thereat, receives mostly fresh ~ ;~
~ ~ :
coolant at the highest impingement ~elocity. Other features ~;
are also present to enhance the capacity of the coolant to
cool the blade.
SUMMARY OF THE INVF.NTION
An integrally cast turbine blade for a gas turbine
engine having a root and a shank portion terminating in a
blade platform and an airfoil portion extending from the
platform to the blade tip is described. A plurality of ;
internal separate serpentine passages permit cooling fluld
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to enter at the root portion and traverse the airfoll por-
tion in predominantly radially extendirlg paths w:lth each
passage exhausting at the blade tip. A separate radially
extending chamber adJacent the leading edge of the airfoil
portion exhausts at both the blade tlp and platform and
receives the cooling fluid from the initial leg of one of
the serpentine passages through a row of small apertures in
a separating wall which provides sufficient velocity to the
coolant to cause impingement against the inner wall of the
leading edge of the chamber for maximum cooling. The exhaust
outlets of this chamber are sized relative to one another
such that maximum cooling occurs approximately 35 to 40~ of
the radial dimension of the airfoil outwardly from the
plakform. The final path of the other serpentine passage,
in addition to exhausting to the tip is in flow communica~
tion with the trailing edge of the airfoil to cool it.
D~SCRIPTION OF THE DRAWINGS
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Figure 1 is an isometric view of the rotor blade ~ -
of the gas turbine blade of the present invention~
Figure 2 is an elevational cross-sectional view
generally along a middle cordal extent of the airfoil,
Figure 3 is a cross-sectional view generally along
line III-III of Figure 2,
Figure 4 is a view similar to Figure 2 of another
embodiment; and,
Figure 5 is a cross-sectional view along V-V of
Figure 4.
DESCRIPTI~N ~F TH:~ PREFE;RRED~MBOPIMENT
~as turbine blades having hol~ow lnterlors defining
pas~ageways for dire~ting coolant therethrough are generally
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well Icno~n in the art. The complete blade can generally be
describecl as comprising, as shown in Figure 1~ a root por-
tion 10, a shank portio~ 12 terminating in a blade platform
14 ~rom which the airfoll portion 16 e!xtends terminating in ~ -
a blade tlp 18. The blade of the present invention is
integrally cast; via a method known as investment casting
wherein a mold, forming the interior passages, is etched or
dissolved after the metal forming the blade ha~ been cast
around it, so that each individual blade requires an indi~
vidual mold forming the coolant passages therein.
The airfoil portion 16 of the blade~ when mounted
in the gas turbine extends radially through the hot motive -~
gas path whlch imparts a driving force thereto. This alr-
foil portlon 16 thus defines, with respect to the gas flow,
a concave pressure surface 20 and a convex suction surface
22 merging at the upstream end in a rounded leading edge 24
and at the downstream end by a trailing edge 26 to form a
generally continuous smooth exterior surface.
A general discussion of design consideratlon ~or a
hollow cooled blade is available in Sawyers Gas ~urbine
Engineering Handbook, Vol. 1 2nd Ed. pages 100 through 107.
It is therein pointed out that as turbine inlet temperatures
exceed 2200F, areas of locally hlgh temperatures would ~ !
occur in the blade that would cause thermal fatigue due to
thermal gradients between these areas and the cooler areas
of the blade. Such temperatures would also cause the metal
alloy of the blade to break down, causing early failure~
Such areas o~ high temperature typlcally are the leading
edge and the traillng edge of the air~oil. To eliminate
such occurrences it is important that these areas receive
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primary cooling consideration.
Thus, rererring to F1i~;ure 2, the coolant ~low
paths through the blade are shown. ~s thereln seen, there
are two separate primary fLow paths 28 and 30, each havlng a
multi-path serpentine configurat:Lon with an inlet 32~ 34,
respectively in the root portion and an outlet 36, 3~,
respectively in the blade tip. The separate legs o~ each
path are primarily radially e~tending with 180 return bends
adJacent the blade tip directing the coolant from the intial
leg 40, 42, respectively, into the return leg 44, 46, respec-
tively and another 180 return bend adjacent the platform
for directing the coolant into the final leg 47, 48 respec-
tively for exiting at the radially outermost termination of
the leg through apertures 50, 52 in the blade tip 18. ;-
The initial or inlet leg and return leg of each
serpentine path are separated by a wall 54, 56 respectively
around which the coolant makes the first 180 bend. However, - ;~
it is seen that adjacent the terminal end of each wall are a
series of openings 58, 60 respectively permitting a portion
of the coolant to pass into the return leg thereby bypassing
the 180 return band of the blade tip. This reduces the
pressure losses that would be present if all the coolant was
forced through the 180 bend which would also tend to remove
more heat from the tip than is necessary to maintain it at
desired temperature, thereby decreasing the ability of the
coolant ~o remove heat from other areas downstream o~ the -~`
tip. Thus~ such bypasses reduce pressure losses and heat
build-up in the coolant.
Further, the final leg of each serpentine passage
contains a plurality of flow interrupters 62g 64 respectively
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extending between opposing walls of the passage ~o that coo-
lant flowing therethrough is alternately accelerated and
decelerated, inducing turbulent flow fc)r greater heat trans- l ~
~er from the internal walls of the passage to the coolant. ~ ;
Also lt should be noted that the inlet 32, 34 of
each separate passage in the root of the blade comprises a
plurality of individual openings separated by walls 70 which
terminate radially in the shank portion where the thlckness ;~
of the blade has increased to insure structural integrity. l~
10 In this area the inlets have smooth wall transition zones 72 ~ -
for directing the coolant into the separate M ow paths with
minimal pressure losses. ~
Still referring to Figure 2 it is seen that a ~;
separate radially extending chamber 74 is provided ad~acent
the leading edge 24 of the air foil 16. Chamber 74 is
separated from the initial leg of the adjacent serpentlne
flow path by a common wall 76. A radially extending row of
small apertures 78 are formed through the wall ~or coolant
flow communication between leg 40 and flow chamber 74 and
are sized such that the velocity of the coolant flowing
therethrough into this chamber causes impingement on the
wall of the chamber closest to the leading edge for impinge~
ment cooling (as discussed in Sawyers) of the critical
leading edge.
Chamber 74 has one exhaust opening 80 at the blade
tip and another exhaust opening 82 directed generally up~
stream ad~acent the blade platform 14. The two exhaust
openings are sized such that approximately 40% of the coolant
i~ exhausted through openin~ 82 ~d~acent the pl~form ~nd
~0% through the blade tip opening 80~ ~urt~er, the openings
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are sized such that the velocity of the flow of the coolant
is sufficiently low so that the coolant entering through the
row of open:lngs 78 ls able to penetrate the outflowing
coolant for impingement on the lnternal wall. Thus with the
exhaust split as above described, it iS axiomatic that at a
point radially inwardl~ from the tlp exhaust a dimenslon
equal to approximately 60~ of the radial dimension between ~ ~;
the two exhaust openings, is a point where the impingement
coolant flow splits such that any coolant entering the
chamber radially above that point exits through the tip
exhaust 80 and coolant enterlng through the openings 78 ;~
below that point exits through the platform exhaust 82.
Thus, coolant at the point of the split will be the least
contaminated (i.e., have minimal mixing with other coolant ;~
that has been exposed to a heated surface) such that it will `~
have the greatest retained cooling ability and will also
have the greatest impingement velocity as it wi71 not have
to flow through outwardly flowing coolant. Any impingement
coolant entering at other points will be mixed, to some
extent, with coolant that has already experienced a temper~
ature rise, such that the further from thls point of split
flow, the greater the temperature of the coolant flowlng
within the chamber that is mixed with the impingement air
whlch, decreases, to some extent, the ability of the mixture
to cool. Thus, the greatest cooling occurs at a point a
distance inwardly from the tip approximately 60% o~ the ~ ;
radial dimension of the airfoil. This point corresponds to
the ~iclnity of the airfoll where a combination o~ motive
~a~ temperature (due to temperature proflling ~8 çxplained
30 in Sawyers, see Figure 49, page 107 of the above identified
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reference) and the centri~ugal force produce the greatest
blade destructive conditions. That i9 s the greate~t temper-
ature after profilln~ generally appears at the midpolnt of
the radial extent of the alr foil whereas the centri~ugal
~orce lncreases from zero at the blade tip to a maximum of
the blade root. Thus, radially inwardly of the midpoint the
temperature o~ t~e hot gas remains extremely high and the
centrifugal force is relatlvely large, And although from
this point lnwardly the centrifugal ~orce continues to rise, !' ~;
the gas temperature of the corresponding point decreases
sufficiently rapldly to reduce the likelihood of blade
failure from the combination of heat and stress,
The trailing edge 26 o~ the airfoil portion 16 ,
also contalns a row of axially extending apertures 84 opening
into the exhaust leg 48 of the second serpenkine path 34 so
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that a metered portion of the coolant passes therethrough ko
cool this portion of the blade with the remainder exiking
through the blade tip opening 52. As is we~l known in the
art, the coolant exiting the blade tlp provldes a seallng ;~
effect betwee~ the blade tip and the ad~acent shroud struc~
ture o~ the turbine so that hot motive ~luid must rlOw over
the airfoil portion of the blade.
Referring now to Figure 3, an axial cross sectlon
to the airfoil portion of the blade shows the leading edge
chamber 74, wall 76 wlth lmpingement apertures 78 3 the
initial leg 40, return leg 44 and exhaust leg 47 of the
first serpentine path 28, with the exhaust leg having a ~low
accelerator 62 thereln. Al~o shown i~ th~ flr~t le~ 42~ ~h~
return leg 46 and the exhaust leg 48 o~ tha seço~d ~erpen~in~
path 30 with trailinæ edge cooling holes 84 through the
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trailing edge. It is hexe emphasi~,ed l;hat, becau~e of the
crlticality of cool:Lng the leading edge, the lmpingement
cooling coolant as obtained from the flrst leg 40 of the ~,
initial path wherein the coolant ls at its highest pressure
and lowest temperature as it passes through the serpentine
path and therefore capable of the greatest cooling capacity,
Another embodiment of the invention shown ln
Figures 4 and ~ describes a coolant circuit through the
blade similar to that heretofore described but with modifi-
cations for increasing the efficiency of the coolant system.
Thus referring to Figure 4 it is seen that rlbs 86protrude from the otherwise smooth walls of each coolant
flow path generally transverse to the direction of flow to -
promote turbulence in the coolant fluid which in turn in~
creases the heat transfer to the coolant from the blade.
Further the trailing edge 26 contains an axial slot 88 `~
extending radially through the airfoil interrupted by a
plurallty of pins 90 extending thereacross so that no single
axial flow channel is defined and the coolant can flow in a
turbulent unconfined manner through the slot to the trailing
edge.
It should be emphasized that the greatest pressure
differential in the coolant clrcuit exist between the inlet
34 of the second serpentine path and the traillng edge 26 of
the airfoil in that the motive gas pressure at this area of
the blade is less than at any other point of coolant fluid
exhaust. Thus, this pressure differential results in a
greater velocity to the coolant, which in turn results in a
greater ability to cool the ad~acent structure. Further,
whereas the velocity in the first embodiment was maintained
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through the apertllres 84 i.n the trai.lirlg edge, in the em-
bodiment of ~igure 4, the flow ls allowed to assume a turbu-
lent pattern of even greater cooling ability. This is :
referred to as pin-fin cool.ing.
Referring now to Figure 5, lt is seen the modifi- .
cations also include a radially extendi.ng row of glll holes
92 between the leading edge impingement chamber 74 and the
suction surface 22 of the airfoil 16. It is expected that
the coolant will exit all gill holes at a uniform flow rate
and that approximately 1/3 to 1/2 of the coolant flow to the
chamber 74 will exit via the gill holes. However3 the
remainlng coolant will exit the tip exhaust 80 and platform
exhaust 82 in the ratio previously described to attain the
same result of maximum cooling in the vicinity of the
greatest likelihood of stress failure of the blade. ~hus, ~ ~
exhaust of the impingement chamber is further utilized to ~ 7 ~ ;
establish surface film cooling on the suction side of the
blade. Other surface film cooling can be provided by a
radial row of outlet openings 94 from the return leg 44 of ;;~
the first passage to the suction side surface 22. Such film
cooling utilizes partially spent coolant to establish a
boundary layer on the surface that reduces the heat load
through the blade due to the hot motive gases. Thus, in the
fi.rst instances, although not cooling the blade, it is
ef~ective to reduce heat transferred to the blade which
ultimately requires less internal coolant to cool the blade,
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and therefore a more efficient engine.
With any or all of these modifications the capacity
: of the coolant to maintain the blade at a reduced acceptable
temperature, although hot motive gases are at a temperature
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well above that whlch would cau~e ~ailure of' the blade, can
b e enhanc ed .
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