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Publication numberUS3891348 A
Publication typeGrant
Publication dateJun 24, 1975
Filing dateApr 24, 1972
Priority dateApr 24, 1972
Also published asDE2320581A1, DE2320581C2
Publication numberUS 3891348 A, US 3891348A, US-A-3891348, US3891348 A, US3891348A
InventorsAuxier Thomas A
Original AssigneeGen Electric
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Turbine blade with increased film cooling
US 3891348 A
Abstract
An air-cooled turbine blade having a plurality of serially spaced cavities therein is provided with cooling air inlets at its base. One of the inlets directs a cooling flow into a cavity proximate the trailing edge of the blade for cooling the blade surfaces defining this cavity. From this cavity, the flow is directed in serpentine fashion into a second cavity remote from the trailing edge for cooling the surfaces defining this latter cavity. The cooling flow is then directed in a film from the second cavity onto the outer surface of the blade for cooling the latter.
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United States Patent 1 1 I 3,891,348 Auxier June 24, 1975 1 1 TURBINE BLADE WITH INCREASED FILM 3,560,107 2/1971 Helms 4l6/90 3,628,885 12/1971 Sidenstick ct a1 416/97 COOLING Primary Examiner-Samuel Feinberg Attorney, Agent, or FirmDana F. Bigelow; Derek P. Lawrence [57] ABSTRACT An aircooled turbine blade having a plurality of serially spaced cavities therein is provided with cooling air inlets at its base. One of the inlets directs a cooling flow into a cavity proximate the trailing edge of the blade for cooling the blade surfaces defining this cavity. From this cavity, the flow is directed in serpentine fashion into a second cavity remote from the trailing edge for cooling the surfaces defining this latter cavity. The cooling flow is then directed in a film from the second cavity onto the outer surface of the blade for cooling the latter.

7 Claims, 5 Drawing Figures TURBINE BLADE WITH INCREASED FILM COOLING The invention herein described was made in the course of or under a contract. or a subcontract thereunder. with the United States Department of the Air Force.

BACKGROUND OF THE INVENTlON This invention relates to blades for use in turbomachinery and. more particularly. to air-coolcd blades of the aforesaid variety.

In turbomachinery. a flow of pressurized working fluid is directed onto a plurality of turbine blades mounted upon rotatable discs for imparting momentum thereto. whereby the kinetic energy of the fluid flow may be transformed into torque. In a number of applications of turbomachine concepts. and particularly with respect to turbojet engines, the working fluid flow is heated to extremely high temperatures. and travels at extremely high velocities. As a result. it has become requisite to discover ways in which to maintain the performance and reliability of turbine blades subjected to such a working environment.

Improvements in metal alloys provided early solutions to the contemporaneous problems of mechanical strength and heat resistance of turbine blades. However. increased demands for performance and ever larger power outputs have mandated blade design variations in addition to improvements in material compositions. Objectives of the design variations have been to provide means for passing cooling fluids to or through portions of the blades subjected to particular heating while reinforcing the structure of turbine blades in the areas of particular mechanical stress.

Members of one variety of blades resulting from the application of these criteria have included a plurality of radially extending cavities serially spaced between the leading and trailing edges of the blades. The cavities perform the function ofdirecting a flow ofcooling fluid through the interior of the turbine blades in order to cool respective portions thereof. Cross ribs extending between adjacent cavities serve to increase the strength of the blades in the directions of stress to which the blades are subjected. ln blades of this variety. it has been the conventional practice to provide cooling air inlets to the cavities which open through the base or tip of each blade to passages in the disc upon which the blades are mounted or to a plenum surrounding the disc. The cooling flow is passed from these inlets through the cavities and eventually is dumped out of the cavities through exits into the environment thereof for expulsion with the working fluid.

In modern. high-powered turbomachinery. overall operating efficiency suffers when turbine blades are cooled by the inefficient application of large amounts of cooling air to the individual blades. since the work required to provide cooling air negates a similar amount of output available from the engine. (onsequcntly. blades using minimum quantities of cooling fluid are desirable. Thus. it has become increasingly important to make full cooling use of the cooling fluid passed through each blade.

Variation of the flow path of the cooling fluid through the blades has been suggested. whereby an increased portion of the available cooling power of the fluid is utilized. A common approach has been to direct a given cooling flow in a serpentine path serially through a number of adjacent cavities prior to the expulsion thereof. It has been further suggested in the prior art to pass a flow of cooling fluid in this serpentine fashion from the trailing edge cavity to one or more serially adjacent cavities before expelling the cooling fluid from apertures near the blade base. tip end and/or trailing edge. While the serpentine path ofthe fluid increases the utilization of available cooling power. the dumping of the used fluid from the tip. base end or trailing edge of the blade fails to comprehend further use to which the cooling fluid might be put. namely utilization of the fluids film cooling potential. This potential has in the past been utilized with respect to cooling flow introduced into leading edge cavities. But the film cooling potential remaining in trailing edge cooling flow upon completion of its serpentine flow path has not been appreciated.

SUMMARY OF THE INVENTION It is therefore an object of the present invention to provide an air-cooled turbine blade having a plurality ofinternal cavities therein with means for further utilizing the cooling potential of a flow of cooling air subsequent to the utilization of this air for the cooling of the blade surfaces defining the internal cavities.

It is a further object of the present invention to accomplish this increased use of cooling potential by directing the flow of cooling air from an internal cavity onto the outer blade surface in a cooling film.

It is a more particular object of the present invention to perform the directing ofthe cooling air as a film onto the outer blade surface after completion by the air of a serpentine flow path through serially adjacent internal cavities originating with the trailing edge cavity.

The present invention seeks to more fully utilize the cooling power of the quantity of cooling fluid supplied to turbine blade trailing edge cavities by expelling particular portions ofthat fluid (which otherwise would be dumped into the gas stream) in a film onto the outer surface of the blade. In order to accomplish this. the present invention provides a number of exit apertures communicating the outer blade surface to the interior of a particular blade cavity. This cavity is the one which the cooling flow reaches upon completion of a predetermined flow path through the blade. The apertures are arranged in a manner appropriate to the formation of a cooling film upon the blades outer surface. The resulting fluid film serves to convectively remove heat from this surface as well as to form a barrier against direct impingement upon the blade by the hot working fluid. As a result. the external blade surface temperature remains lower. so that less cooling air need be provided. Consequently. the overall efficiency of the engine is enhanced.

Further objects of the present invention will become apparent from the detailed description of a preferred embodimentcontained hereinafter as illustrated by the following figures wherein:

FIG. I is a section view of a typical turbojet engine showing the essential elements thereof;

FIG. 2 is a partial section view of the turbojet engine of FIG. 1 showing the turbine arrangement in greater detail;

FIG. 3 is a section view of the blade of FIG. 2 taken along lines 3-3;

FIG. 4 is a half-section view of the turbine blade of the present invention; and

FIG. 5 is a schematic diagram ofthe flowpath ofcooling air within the turbine blade of FIG. 2.

DESCRIPTION OF A PREFERRED EMBODIMENT The following description amplifies the particular embodiment of the present invention depicted in the accompanying drawings. ()ne skilled in the art might easily recognize numerous changes which may be made in the structure of this embodiment without departing from the spirit of the present invention.

The turbojet engine depicted in FIG. I comprises the basic elements of typical machinery of this variety. A substantially cylindrical housing surrounds a compressor I0. combustors I]. and a turbine 12, all disposed about rotatable shaft 13. As is well known in the art. atmospheric air enters the machine from the left to be pressurized. heated and expelled to the right to provide usable thrust. More particularly. air enters from the left and is operated upon by the compressor (in combi nation with the shape of the lead end of shaft 13) to be pressurized and directed. in part. into combustors 11. Heat energy is added to the air within the combustors by the burning of appropriate fuel supplied thereto. Working fluid. which is a combination of air and burned fuel. exits at the right end of the combustors and impinges the plurality ofturbine blades 14 carried by a number of adjacent discs making up the turbine l2. The impingement of the turbine blades 14 by the working fluid serves to drive the turbine in rotation. which rotation is imparted to shaft l3. The rotation of shaft 13 is the motivating force for the operation of the compressor 10 at the forward end of the machine.

Turbine blades 14 must be extremely strong and heat resistant in order to withstand the force and heat of the impinging working fluid. The cooling system of the present invention. by which the blades l4 are protected from overheating. is depicted in H05. 2 through 5 and operates by making use ofa portion of the air operated upon by compressor 10 but not directed into the combustors 1].

H6. 2 shows a typical turbine blade 14 and its cooperation with shaft 13 and elements of the cooling system which are elucidated hereinafter. Blade 14 includes a blade shell [6 in the shape of an airfoil, and a platform [8 adapted to cooperate with disc 19 by which the plurality of blades are supported. Blade shell 16 has an outer surface 17 and a plurality ofinner cavities which will also be discussed hereinafter. At the blade tip (the blade end opposite platform 18) is a closure 20 which separates the blade inner cavities from environmental atmosphere. and which may be integral with the blade shell 16 or a separate piece affixed thereto. The blade shell l6 further has a leading edge 26 and a trailing edge 28.

In the embodiment ofthe invention disclosed in FIG. 2, an aperture 22 in shaft [3 permits the passage therethrough of cooling air from a cooling air expander 24 cooperating with an appropriate plenum (not shown) for the delivery of cooling air to the blade. It is recognized that the present invention is equally applicable to blades which are designed to be cooled by the application of air provided through inlets located near the tip of the blade rather than near the platform or by internal turbine circuits that do not use a cooling air expander. The present embodiment is to serve only as an example and not be considered the only embodiment of the present in ention.

The cross-sectional viewsof blade [4 depicted in HUS. 3 and 4. taken together with the schematic flow path of Fl(]. 5. disclose the blade structure which directs the application ofthe cooling air fed to the blade platform 18 as described above. FIGS. 3 and 4 show that the turbine blade ofthe present embodiment of the invention incorporates first. second and third cavities labeled 30. 32 and 34. respectively. which are disposed serially adjacent to one another between leading edge 26 and trailing edge 28 of the blade shell lb. The cavities are defined within the shell by inner shell surfaces 3|. 33 and 35. respectively. The three cavities are of shape and size determined to be appropriate for the optimization of cooling efficiency and mechanical blade strength. Cavity 30 is disposed proximate trailing edge 28 of the blade. while cavity 32 is disposed remote from the trailing edge 28. Cavity 34 is disposed proximate the leading edge of the blade.

lnserts 36 and 38 are disposed respectively within cavities 32 and 34. and respectively bear pluralities of orifices 37 and 39 for the distribution of cooling air in an impinging flow against the inner surfaces of each respective cavity. The inner surfaces 3| of cavity 30 are provided with a plurality of protrusions 31a appropriately positioned to enhance the turbulence of flow for minimum pressure drop. As defined by the state of the art. the impinging flow associated with cavities 32 and 34 and the turbulent flow associated with cavity 30 are superior in cooling characteristics to the flows which would occur within the cavities absent the provisions described.

The blade 14 of the present embodiment is provided with two inlets 40 and 42 for the entry of cooling fluid from passage 22 (see FIG. 2). It is noted that the two inlets 40 and 42 service cavities 30 and 34, respectively. Means for passing cooling fluid to cavity 32 in' eludes a passage 44 between cavities 32 and 30 disposed in proximity to the tip of blade 14 and remote from inlet 40. The two cavities 30 and 32 and passage 44 define a serpentine path for the cooling fluid entering inlet 40.

The definition of appropriate exits for the cooling fluid. which exits combine with the foregoing blade structure to maximize the utilization of the cooling fluid. accomplishes the objects of the present invention. As in the prior art. a minor portion of the cooling fluid which has been introduced through trailing edge inlet 40 is exhausted through a plurality of exit apertures 46 provided at the trailing edge 28 of the blade shell 16. As is also prevalent in the prior art. a portion of the air entering leading edge cavity 34 through inlet 42 is exhausted through three groups of exit apertures 48. 50 and 52 positioned and adapted to direct the exit ing fluid in a film across various portions of the outer surface of the blade shell in.

Film cooling has been found to be useful to increase the use to which cooling air may be put. whose cooling potential has not been exhausted during application to the inner surfaces of the blade shell. lf. after passing from the inlet 42 and through orifices 39 in insert 38 and against surfaces 35. the air within cavity 34 re mains at a temperature lower than that existing in the working fluid near the outer surfaces of the blade shell 16. the passing of this air out of cavity 34 in a film across such outer surfaces would serve to cool them and thus to make further use of the cooling flow. it is precisely to this film-cooling concept that the present invention is directed While the prior art has comprehended the use of exit film cooling to maximive the utili/ation of the cooling power of air fed into leading edge cavities. it has been common practice with respect to cooling flow fed into trailing edge cavities to dump the flow out of exit apertures at the tip. base or trailing edge of the blade after an internal serpentine path has been completed. Wherever the working fluid near the outer surfaces of the blade to which this latter flovv might be directed is at a temperature higher than the exiting cooling flow. this practice constitutes a waste of cooling power.

The present invention thus provides a plurality of spaced exit apertures 54 through which the cooling fluid may be directed onto the outer surface of blade shell l6 downstream of apertures 54. Apertures 54 are located substantially along a radial line between the ends of blade shell 16 and are configured appropriately for the formation of a cooling film upon the outer blade surface. The film thus formed serves as a barrier to protect the blade from the direct impingement of the hot working fluid. Further. the film serves to remove heat from the blade surface by convective heat flow. This added usage of the cooling power of the cooling fluid allows the turbine blade to be cooled to the same extent as previously. but with the expenditure of less cooling fluid. As described above. beneficial effects upon the overall efficiency of the turbomachine are thus achieved.

The operation of the cooling system of the present invention will now be described with the aid of the alphabetical designations of locations depicted in FIGS. 3 and 4 and represented schematically in FIG. 5. Cooling air from the plenum is passed through air passage 22 of FIG. 2 to the platform 18 of blade l4 and into inlets 40 and 42 of FIG. 4. That portion of the flow entering inlet 42 passes from point A below cavity 34 to point B within cavity 34 and into contact with insert 38. The air is passed through orifices 39 and into the area represented by point C. which is defined by the insert 38 and inner surfaces 35 of the blade shell. The impingement of this air against surfaces 35 serves to cool these surfaces before the air is exhausted through exit apertures 48 and St) to points D and E. respectively. The working fluid flowing past points D and E impinges the exiting cooling fluid and. due to the viscous forces therehetween. creates films to the downstream sides of each point which films serve to cool the external surfaces of the blade shell [6 until the films are separated therefrom by turbulence.

The second portion of cooling air flowing into the blade (through inlet 40) passes from point F below the blade shell to point (i proximate the blade tip closure 20. Since the flow enters the blade at the trailing edge at its coolest temperature. a minimum portion of this fluid is forced out of cavity 30 through trailing edge exit apertures 46. The predominant portion of the fluid passes through passage 44 from point 0 to point H within cavity 32. This flow continues past point I within cavity 32 and to point J. While within cavity 30, the cooling fluid acts in turbulent flow to cool surfaces 3!. Having passed into cavity 32. the fluid is directed by orifices 37 of insert 36 into the area represented by point K and into impingement with surfaces 33 for the cooling thereof. Having progressed to this point and having htl been raised in temperature by contact with surfaces 3] and 33 of the blade shell. the cooling fluid remains at a temperature below the external surface temperature of the blade shell. Thus. the fluid still possesses usable cooling power. Accordingly. there are provided apertures 54 through which the fluid is subsequently passed to point l. outside of the blade shell. The viscous forces of the passing working fluid act upon the exiting cooling fluid to create a cooling fllm downstream ofexit orifices 54 upon the outer surface of the blade shell proxi mate trailing edge 28. This film forms a barrier between the outer blade surface and the working fluid. The film also cools the blade surface by convective heat transfer. Consequently. the film serves to further cool the blade by increasing heat transfer to the cooling fluid after its exit from the blade.

In this way. the present invention increases the utilization of the cooling power ofa given quantity of cooling fluid by maximizing the contact with various turbine blade shell surfaces of that portion of the cooling fluid fed into trailing edge cavities. Were the fluid entering inlet 40 to be dumped into the passing working fluid from either the tip or base ends of the blade shell, no film would be created upon the outer surface of the blade by this fluid. and the remaining cooling power thereof would be wasted. By the application of the present invention to turbine blades. reductions in the amount of cooling fluid required to be fed to the rear or trailing edge blade cavities may be effected. with attendant increases in machine efficiency.

While the present invention has been described in conjunction with a preferred embodiment thereof. it is apparent that numerous variations in the application thereof may be made without departing from the spirit of the invention. For example. a turbine blade having a plurality of cavities in a number of larger than the three disclosed might be devised wherein cooling air fed to one of the cavities proximate the trailing edge is passed in a serpentine path through serially adjacent cavities and finally exhausted in a cooling film upon external blade surfaces. Another variation, which was discussed briefly above. might involve the application of cooling fluid through inlets located near the tip rather than the platform of the blade shell. Additionally. while the embodiment herein disclosed passes a cooling film over only one side of the outer surface of the blade shell. a plurality of apertures could easily be applied which would serve to communicate the other side of the shell to a trailing edge cavity for the passing of cooling fluid thercacross.

Other modifications of the described embodiment of the invention will occur to those skilled in the art within the scope of the present inventive concept without departing from the spirit thereof.

What is claimed as new and desired to be secured by Letters Patent of the United States is:

I. In a turbomachine. an air-cooled blade comprising:

a blade shell having an outer surface. first. second and third inner surfaces. and a leading edge and a trailing edge;

a first cavity in said shell proximate said trailing edge and formed by said first inner surface of said shell.

a second cavity in said shell remote from said trailing edge and formed by said second inner surface of said shell;

a third cavity in said shell proximate said leading edge and formed by said third inner surface of said shell;

means for introducing cooling air into said first cavity for cooling said first inner surface:

means for introducing cooling air into said third cavity for cooling said third inner surface:

a plurality of passageways formed through said trail ing edge and communicating with said first cavity for efflux of a first portion of said first cavity cooling air therefrom;

means for directing the remainder of said first cavity cooling air from said first cavity into said second cavity for cooling said second inner surface: and

means for directing all of said cooling air delivered to said second cavity in a film from said second cavity onto said outer surface for cooling said outer surfacev 2. The blade of claim I further comprising first and second blade ends and a pair of ribs located internally of said shell and extending between said first and second blade ends. wherein said means for introducing cooling air into said first cavity is disposed proximate said first blade end. and said means for directing said cooling air into said second cavity comprises an opening in one of said ribs disposed proximate said second blade end.

3. The blade of claim I further comprising a radially outward tip end and a radially inward platform wherein said means for directing said cooling air onto said outer surface comprises a plurality of spaced apertures providing communication between said second cavity and said outer surface. disposed between said tip end and said platform and substantially aligned along a radial line along said blade shell.

4. The blade of claim 3 further including a hollow impingement insert positioned within said second cavity. and said cooling air directing means directs said remainder of said cooling air initially to the interior of said insert.

5. The blade of claim 4 further including a second hollow impingement insert positioned within said third cavity.

6. An air-cooled blade for use in a turbomachinc.

said blade comprising;

a blade shell in the shape of an airfoil having a leading edge and a trailing edge and further having an outer surface and first. second and third inner surfaces and first and second ends;

a blade platform proximate the first end of said shell;

first. second and third serially spaced cavities defined respectively by said first. second and third inner surfaces of said shell and disposed respectively proximate said trailing edge. remote from said trailing edge. and proximate said leading edge;

a blade closure proximate the second end of said shell for separating said cavities from an environmental atmosphere:

means for introducing a first flow of cooling air through said platform into said first cavity for cooling said first inner surface;

a plurality of passageways formed through said trailing edge and communicating with said first cavity for efflux of a first portion of said first cavity cooling air therefrom.

means proximate said closure for directing the remainder of said first flow from said first cavity into said second cavity for cooling said second inner surface;

means for directing all of said first flow delivered to said second cavity in a film from said second cavity onto said outer surface for cooling said outer surface:

means for introducing a second flow of cooling air through said platform into said third cavity for cooling said third inner surface; and

means for directing said second flow in a film from said third cavity onto said outer surface for further cooling said outer surface. l

7. The blade of claim 6 wherein said means for directing said first flow onto said outer surface comprises a plurality of spaced apertures substantially aligned along a radial line between said first and second ends of said blade shell.

Patent Citations
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US3301526 *Dec 22, 1964Jan 31, 1967United Aircraft CorpStacked-wafer turbine vane or blade
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Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US4168938 *Jan 28, 1977Sep 25, 1979Rolls-Royce LimitedBlade or vane for a gas turbine engine
US4221539 *Apr 20, 1977Sep 9, 1980The Garrett CorporationLaminated airfoil and method for turbomachinery
US4257737 *Jul 10, 1978Mar 24, 1981United Technologies CorporationCooled rotor blade
US4507051 *Nov 10, 1982Mar 26, 1985S.N.E.C.M.A.Gas turbine blade with chamber for circulation of cooling fluid and process for its manufacture
US4645415 *Aug 11, 1986Feb 24, 1987United Technologies CorporationAir cooler for providing buffer air to a bearing compartment
US4705455 *Dec 23, 1985Nov 10, 1987United Technologies CorporationConvergent-divergent film coolant passage
US5120192 *Mar 12, 1990Jun 9, 1992Kabushiki Kaisha ToshibaCooled turbine blade and combined cycle power plant having gas turbine with this cooled turbine blade
US5156526 *Dec 18, 1990Oct 20, 1992General Electric CompanyRotation enhanced rotor blade cooling using a single row of coolant passageways
US5165852 *Dec 18, 1990Nov 24, 1992General Electric CompanyRotation enhanced rotor blade cooling using a double row of coolant passageways
US5503529 *Dec 8, 1994Apr 2, 1996General Electric CompanyTurbine blade having angled ejection slot
US5591002 *Aug 1, 1995Jan 7, 1997General Electric Co.Closed or open air cooling circuits for nozzle segments with wheelspace purge
US5660524 *Jul 13, 1992Aug 26, 1997General Electric CompanyAirfoil blade having a serpentine cooling circuit and impingement cooling
US6183192Mar 22, 1999Feb 6, 2001General Electric CompanyDurable turbine nozzle
US6193465 *Sep 28, 1998Feb 27, 2001General Electric CompanyTrapped insert turbine airfoil
US6283708Dec 3, 1999Sep 4, 2001United Technologies CorporationCoolable vane or blade for a turbomachine
US9133717 *Jan 8, 2009Sep 15, 2015Ihi CorporationCooling structure of turbine airfoil
US20110027102 *Jan 8, 2009Feb 3, 2011Ihi CorporationCooling structure of turbine airfoil
USRE39479 *Feb 5, 2003Jan 23, 2007General Electric CompanyDurable turbine nozzle
EP0151918A2 *Jan 2, 1985Aug 21, 1985National Aeronautics And Space AdministrationMethod and apparatus for cooling high temperature structures with a fluid coolant
EP0151918A3 *Jan 2, 1985Mar 11, 1987National Aeronautics And Space AdministrationMethod and apparatus for cooling high temperature structures with a fluid coolant
Classifications
U.S. Classification416/97.00R, 416/96.00A
International ClassificationF01D5/18
Cooperative ClassificationF01D5/189
European ClassificationF01D5/18G2C