CA1087527A - Cooled gas turbine blade - Google Patents
Cooled gas turbine bladeInfo
- Publication number
- CA1087527A CA1087527A CA294,987A CA294987A CA1087527A CA 1087527 A CA1087527 A CA 1087527A CA 294987 A CA294987 A CA 294987A CA 1087527 A CA1087527 A CA 1087527A
- Authority
- CA
- Canada
- Prior art keywords
- coolant
- blade
- chamber
- path
- leading edge
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Expired
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Abstract
COOLED GAS TURBINE BLADE
ABSTRACT OF THE DISCLOSURE
A cooled gas turbine blade having two separate serpentine coolant flow paths, each having an entry in the blade root and an exhaust in the blade tip. The initial leg of one path is in flow communication with a chamber adjacent the leading edge of the blade through a plurality-of small openings along the radial extent of the airfoil portion for impingement cooling the leading edge with relatively high pressure, low temperature coolant. The chamber is exhausted through a blade tip opening and a platform opening which are related in size to maximize impingement cooling affect at a predetermined radial portion of the leading edge. The flow paths are configured to minimize pressure drop therein and enhance the effectiveness of the coolant.
ABSTRACT OF THE DISCLOSURE
A cooled gas turbine blade having two separate serpentine coolant flow paths, each having an entry in the blade root and an exhaust in the blade tip. The initial leg of one path is in flow communication with a chamber adjacent the leading edge of the blade through a plurality-of small openings along the radial extent of the airfoil portion for impingement cooling the leading edge with relatively high pressure, low temperature coolant. The chamber is exhausted through a blade tip opening and a platform opening which are related in size to maximize impingement cooling affect at a predetermined radial portion of the leading edge. The flow paths are configured to minimize pressure drop therein and enhance the effectiveness of the coolant.
Description
RAc~GRotJMn or~ T~ INV~h~TION
.
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 .
~: :
~7~
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 '~' ~75~
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
~ .
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 .
~37~27 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 S;27 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 ~- . .
52~
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 -6~
752~7 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 ~7- ~
' ~,~-~''.' 75~
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 ~ .
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 ., . .
75~
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 _9~
, ~.
gl(3~752~
. . .
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, ,:
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 -1 0~
well above that whlch would cau~e ~ailure of' the blade, can b e enhanc ed .
'''~ '' .
-`'
.
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 .
~: :
~7~
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 '~' ~75~
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
~ .
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 .
~37~27 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 S;27 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 ~- . .
52~
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 -6~
752~7 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 ~7- ~
' ~,~-~''.' 75~
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 ~ .
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 ., . .
75~
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 _9~
, ~.
gl(3~752~
. . .
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, ,:
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 -1 0~
well above that whlch would cau~e ~ailure of' the blade, can b e enhanc ed .
'''~ '' .
-`'
Claims (6)
1. A cooled gas turbine blade having a root and shank portion terminating in a blade platform and an airfoil portion extending radially outwardly therefrom to the blade tip, said airfoil portion defining a leading edge and trailing edge interconnected by a concave pressure surface and a convex suction surface, a plurality of separate internal coolant passages for directing coolant through the blade to maintain the blade at a temperature less than the temperature of the motive gas in the turbine, said plurality of passages comprising:
first and second separate passages, each of said passages having an inlet in said root portion and an outlet in said airfoil portion, said inlet and outlet of each passage connected through a serpentine configuration of substantially radially extending paths including an initial path and a final path and wherein said paths are connected for flow communi-cation by at least one intermediate radial path and return bends generally adjacent the tip and shank portions;
a radially extending chamber generally adjacent the leading edge of said airfoil portion and having a first outlet adjacent the blade tip and a second outlet adjacent the blade platform;
a wall member separating the initial radially extending path of said first separate passage and said chamber; and, a plurality of openings extending through said wall member over most of its radial extent to permit coolant flow communication between said last-named path and said chamber, said coolant impinging on the internal wall of said chamber adjacent said leading edge to have a major effect in cooling said leading edge and subsequently exiting said chamber through said first and second outlets;
the trailing edge of said airfoil portion defining collant flow passages from the final radial path of said second separate passage to exhaust at said trailing edge generally throughout the radial extent of said airfoil portion; and wherein said first and second outlets from said chamber are sized relative to each other to meter a prede-termined ratio of coolant to be exhausted from each outlet in accordance with the optimal radial position of the leading edge for receiving the greatest cooling effect from impingement cooling.
first and second separate passages, each of said passages having an inlet in said root portion and an outlet in said airfoil portion, said inlet and outlet of each passage connected through a serpentine configuration of substantially radially extending paths including an initial path and a final path and wherein said paths are connected for flow communi-cation by at least one intermediate radial path and return bends generally adjacent the tip and shank portions;
a radially extending chamber generally adjacent the leading edge of said airfoil portion and having a first outlet adjacent the blade tip and a second outlet adjacent the blade platform;
a wall member separating the initial radially extending path of said first separate passage and said chamber; and, a plurality of openings extending through said wall member over most of its radial extent to permit coolant flow communication between said last-named path and said chamber, said coolant impinging on the internal wall of said chamber adjacent said leading edge to have a major effect in cooling said leading edge and subsequently exiting said chamber through said first and second outlets;
the trailing edge of said airfoil portion defining collant flow passages from the final radial path of said second separate passage to exhaust at said trailing edge generally throughout the radial extent of said airfoil portion; and wherein said first and second outlets from said chamber are sized relative to each other to meter a prede-termined ratio of coolant to be exhausted from each outlet in accordance with the optimal radial position of the leading edge for receiving the greatest cooling effect from impingement cooling.
2. Structure according to claim 1 wherein approxi-mately 40% of the coolant exiting said chamber through said outlets passes through said second outlet so that the radial position of maximum impingement cooling is at a point generally 40% of the distance between said first and second outlets from said second outlet.
3. Structure according to claim 2, wherein the final radial path of each of said separate passages contains rib members projecting thereinto to induce turbulent flow in the coolant and thereby enhance the cooling effect.
4. Structure according to claim 3, wherein the inlet to each separate passage in said root portion com-prises a plurality of individual openings leading to a radially outwardly enlarged transition zone, said openings being separated by walls providing strength to said root portion.
5. Structure according to claim 3, wherein a radially extending row of apertures connect said chamber with the suction surface of said airfoil adjacent said leading edge to deliver a film of coolant to said suction surface.
6. Structure according to claim 1, wherein the initial and at least one intermediate path of each of said separate passages are separated by a wall member having an aperture therein upstream of the return bend joining said initial and intermediate path to permit a portion of said coolant flow to enter said intermediate path therethrough and bypassing said return bend.
Applications Claiming Priority (2)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US76724577A | 1977-02-10 | 1977-02-10 | |
| US767,245 | 1977-02-10 |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA1087527A true CA1087527A (en) | 1980-10-14 |
Family
ID=25078924
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA294,987A Expired CA1087527A (en) | 1977-02-10 | 1978-01-16 | Cooled gas turbine blade |
Country Status (4)
| Country | Link |
|---|---|
| JP (1) | JPS5399116A (en) |
| AR (1) | AR212148A1 (en) |
| CA (1) | CA1087527A (en) |
| IT (1) | IT1092591B (en) |
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4604031A (en) * | 1984-10-04 | 1986-08-05 | Rolls-Royce Limited | Hollow fluid cooled turbine blades |
| US4992026A (en) * | 1986-03-31 | 1991-02-12 | Kabushiki Kaisha Toshiba | Gas turbine blade |
Families Citing this family (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| JPS58170801A (en) * | 1982-03-31 | 1983-10-07 | Toshiba Corp | Blade for turbine |
| JPS5918203A (en) * | 1982-07-21 | 1984-01-30 | Agency Of Ind Science & Technol | Blade of gas turbine |
| JPS5931273U (en) * | 1982-08-20 | 1984-02-27 | カシオ計算機株式会社 | Stand mechanism for small electronic devices |
| US6273682B1 (en) * | 1999-08-23 | 2001-08-14 | General Electric Company | Turbine blade with preferentially-cooled trailing edge pressure wall |
| US9528379B2 (en) * | 2013-10-23 | 2016-12-27 | General Electric Company | Turbine bucket having serpentine core |
Family Cites Families (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB1188401A (en) * | 1966-02-26 | 1970-04-15 | Gen Electric | Cooled Vane Structure for High Temperature Turbines |
| JPS5138372B2 (en) * | 1973-04-07 | 1976-10-21 |
-
1978
- 1978-01-16 CA CA294,987A patent/CA1087527A/en not_active Expired
- 1978-01-31 AR AR27091378A patent/AR212148A1/en active
- 1978-02-09 IT IT20103/78A patent/IT1092591B/en active
- 1978-02-10 JP JP1365578A patent/JPS5399116A/en active Pending
Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US4604031A (en) * | 1984-10-04 | 1986-08-05 | Rolls-Royce Limited | Hollow fluid cooled turbine blades |
| US4992026A (en) * | 1986-03-31 | 1991-02-12 | Kabushiki Kaisha Toshiba | Gas turbine blade |
Also Published As
| Publication number | Publication date |
|---|---|
| AR212148A1 (en) | 1978-05-15 |
| IT1092591B (en) | 1985-07-12 |
| IT7820103A0 (en) | 1978-02-09 |
| JPS5399116A (en) | 1978-08-30 |
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