System and method for airfoil film cooling

Information

  • Patent Grant
  • 6629817
  • Patent Number
    6,629,817
  • Date Filed
    Thursday, July 5, 2001
    23 years ago
  • Date Issued
    Tuesday, October 7, 2003
    21 years ago
Abstract
An airfoil for a gas turbine engine including an inflection that facilitates enhancing film cooling of the airfoil, without adversely affecting aerodynamic efficiency of airfoil is described. The airfoil includes a generally concave first sidewall and a generally convex second sidewall joined at a leading edge and at a trailing edge of the airfoil. A plurality of cooling openings extend between an internal cooling chamber and an external surface of the first sidewall. One cooling opening extends from the cooling chamber into the inflection at a relatively shallow injection angle with respect the airfoil external surface.
Description




BACKGROUND OF THE INVENTION




This application relates generally to gas turbine engines and, more particularly, to methods and apparatus for cooling airfoils used within gas turbine engines.




At least some known gas turbine engines include a compressor, a combustor, and a turbine. Airflow entering the compressor is compressed and directed to the combustor where it is mixed with fuel and ignited, producing hot combustion gases used to drive the turbine. Because components within the turbine are exposed to hot combustion gases, cooling air is routed to the airfoils and blades.




For example, a turbine vane or rotor blade typically includes a hollow airfoil, the outside of which is exposed to the hot combustion gases, and the inside of which is supplied with cooling fluid, which is typically compressed air. The airfoil includes leading and trailing edges, a pressure side, and a suction side. The pressure and suction sides connect at the airfoil leading and trailing edges, and span radially between an airfoil root and an airfoil tip. Film cooling holes extend between a cooling chamber defined within the airfoil and an outer surface of the airfoil. The cooling holes route cooling fluid from the cooling chamber to the outside of the airfoil for film cooling the airfoil. The film cooling holes discharge cooling fluid at an injection angle that is measured with respect to the outer surface of the airfoil.




Because of the curvature distribution of the outer surface of the airfoil between the leading and trailing edges, the injection angles of the cooling holes are typically between 25 and 40 degrees. Cooling fluid discharged from cooling holes having increased injection angles may separate from the surface of the airfoil and mix with the hot combustion gases. Such separation decreases an effectiveness of the film cooling and increases aerodynamic mixing losses.




To facilitate reducing aerodynamic mixing losses, at least some known airfoils include curved film cooling openings. The curved film cooling openings have injection angles as low as 16.5 degrees. However, the cooling fluid may separate from an inner wall of the cooling opening and be discharged in an erratic manner. Furthermore, manufacturing such curved openings is a complex and costly procedure.




BRIEF SUMMARY OF THE INVENTION




In an exemplary embodiment, an airfoil for a gas turbine engine includes an inflection that facilitates enhancing film cooling of the airfoil, without adversely impacting aerodynamic efficiency of the airfoil. The airfoil includes a generally concave first sidewall and a generally convex second sidewall. The sidewalls are joined at a leading edge and at a chordwise spaced trailing edge of the airfoil that is downstream from leading edge. A cooling chamber is defined within the sidewalls, and a plurality of cooling openings extend between the cooling chamber and an external surface of the first sidewall. At least one of the cooling openings extends from the cooling chamber into the inflection at an injection angle measured with respect to an external surface of the airfoil.




In another aspect, a gas turbine engine including a plurality of airfoils that each include a leading edge, a trailing edge, a first sidewall having an outer surface, and a second sidewall having an outer surface is provided. The airfoil first and second sidewalls are connected chordwise at the leading and trailing edges. The first and second sidewalls extend radially from an airfoil root to an airfoil tip, and at least one of the first sidewall and said second sidewall also includes an inflection.




In a further aspect, a method for contouring an airfoil for a gas turbine engine to facilitate improving film cooling effectiveness of the airfoil is provided. The airfoil includes a leading edge, a trailing edge, a first sidewall, and a second sidewall. The first and second sidewalls are connected chordwise at the leading and trailing edges to define a cavity, and extend radially between an airfoil root and an airfoil tip. The method includes the steps of forming an inflection in an outer surface of at least one of the airfoil first sidewall and the airfoil second sidewall, such that the inflection extends a distance radially between the airfoil root and the airfoil tip, and forming at least one opening within the inflection for receiving cooling fluid therethrough from the airfoil cavity to the airfoil outer surface.











BRIEF DESCRIPTION OF THE DRAWINGS





FIG. 1

is schematic illustration of a gas turbine engine;





FIG. 2

is a cross sectional view of a known airfoil that may be used with the gas turbine engine shown in

FIG. 1

;





FIG. 3

is a cross sectional view of an airfoil that may be used with the gas turbine engine shown in

FIG. 1

;





FIG. 4

is a partial cross sectional view of an alternative embodiment of an airfoil that may be used with the gas turbine engine shown in

FIG. 1

; and





FIG. 5

is a cross sectional view of a further alternative embodiment of an airfoil that may be used with the gas turbine engine shown in FIG.


1


.











DETAILED DESCRIPTION OF THE INVENTION





FIG. 1

is a schematic illustration of a gas turbine engine


10


including a fan assembly


12


, a high pressure compressor


14


, and a combustor


16


. Engine


10


also includes a high pressure turbine


18


, and a low pressure turbine


20


. Engine


10


has an intake side


28


and an exhaust side


30


. In one embodiment, engine


10


is a CFM 56 engine commercially available from General Electric Corporation, Cincinnati, Ohio.




In operation, air flows through fan assembly


12


and compressed air is supplied to high pressure compressor


14


. The highly compressed air is delivered to combustor


16


. Airflow (not shown in

FIG. 1

) from combustor


16


drives turbines


18


and


20


, and turbine


20


drives fan assembly


12


.





FIG. 2

is a cross sectional view of a known airfoil


31


including a leading edge


32


and a chord-wise spaced trailing edge


34


that is downstream from leading edge


32


. Airfoil


31


is hollow and includes a first sidewall


36


and a second sidewall


38


. First sidewall


36


is generally convex and defines a suction side of airfoil


31


, and second sidewall


38


is generally concave and defines a pressure side of airfoil


31


. Sidewalls


36


and


38


are joined at airfoil leading and trailing edges


32


and


34


. More specifically, first sidewall


36


is curved and aerodynamically contoured to join with second sidewall


38


at leading edge


32


.





FIG. 3

is a cross sectional view of an airfoil


40


that may be used with a gas turbine engine, such as engine


10


, shown in FIG.


1


. In one embodiment, airfoil


40


is used within a plurality of rotor blades (not shown) that form a high pressure turbine rotor blade stage (not shown) of the gas turbine engine. In another embodiment, airfoil


40


is used within a plurality of turbine vanes (not shown) used to direct a portion of a gas flow path from a combustor, such as combustor


16


, shown in

FIG. 1

, onto annular rows of rotor blades.




Airfoil


40


is hollow and includes a first sidewall


44


and a second sidewall


46


. First sidewall


44


is generally convex and defines a suction side of airfoil


40


, and second sidewall


46


is generally concave and defines a pressure side of airfoil


40


. Sidewalls


44


and


46


are joined at a leading edge


48


and at a chordwise spaced trailing edge


50


of airfoil


40


that is downstream from leading edge


48


.




First and second sidewalls


44


and


46


, respectively, extend longitudinally or radially outward to span from an airfoil root (not shown) to an airfoil tip (not shown) which defines a radially outer boundary of an internal cooling chamber


58


. Cooling chamber


58


is further defined within airfoil


40


between sidewalls


44


and


46


. Internal cooling of airfoils


40


is known in the art. In one embodiment, cooling chamber


58


includes a serpentine passage (not shown) cooled with compressor bleed air.




First and second sidewalls


44


and


46


, respectively, each have a relatively continuous arc of curvature between airfoil leading and trailing edges


48


and


50


, respectively. Additionally, each sidewall


44


and


46


, includes an outer surface


60


and


62


, respectively, and an inner surface


64


and


66


, respectively. Each sidewall inner surface


64


and


66


is adjacent to cooling chamber


58


.




Airfoil


40


also includes an inflection or an area of localized surface contouring


70


. More specifically, near airfoil leading edge region


48


, sidewall


44


is contoured to form inflection


70


, such that a thickness


72


of sidewall


44


remains substantially constant through inflection


70


. In an alternative embodiment, either sidewall


44


or


46


, or both sidewalls


44


and


46


, are contoured to form inflection


70


. In a further embodiment, sidewall thickness'


72


and


74


are variable through inflection


70


. Inflection


70


extends substantially longitudinally or radially between the airfoil root and the airfoil tip.




A plurality of cooling openings


80


extend between cooling chamber


58


and airfoil outer surfaces


60


and


62


to connect cooling chamber


58


in flow communication with airfoil outer surfaces


60


and


62


. In one embodiment, each cooling opening


80


has a substantially circular diameter. Cooling openings


80


discharge cooling fluid through fluid paths known as injection jets. Alternatively, each cooling opening


80


is non-circular. At least one cooling opening


82


extends between airfoil outer surface


60


and cooling chamber


58


within inflection


70


. More specifically, inflection cooling opening


82


has a centerline


84


, and extends through sidewall


44


at an injection angle Ø. Injection angle Ø is formed by an intersection of centerline


84


and a line


86


that is tangent to airfoil outer surface


60


at a point where cooling opening


82


intersects airfoil outer surface


60


. In one embodiment, injection angle Ø is less than approximately 16 degrees.




During operation, although the curvature of airfoil sidewalls


44


and


46


is advantageous in directing combustion gases, contact with the combustion gases increases a temperature of airfoils


40


. To facilitate cooling airfoil


40


, cooling fluid is routed through cooling openings


80


and used in film cooling airfoil outer surfaces


60


and


62


. The injection of cooling fluid into a boundary layer, known as film cooling, produces an insulating layer or film between airfoil outer surfaces


60


and


62


, and the hot combustion gases flowing past airfoil


40


.




Because airfoil inflection


70


permits cooling fluid to be provided to airfoil outer surface


60


through inflection cooling opening


82


at a relatively shallow injection angle Ø, a reduction in coolant injection jet separation is facilitated, therefore enhancing film cooling effectiveness. Furthermore, because inflection


70


facilitates enhancing film cooling effectiveness, reduced amounts of cooling fluid for a set amount of heat transfer may be utilized. Alternatively, because inflection


70


facilitates enhancing film cooling effectiveness, a useful life of airfoil


40


may be facilitated to be extended. Furthermore, aerodynamic losses associated with inflection


70


are facilitated to be reduced because inflection cooling opening


82


injects cooling fluid at a shallow injection angle Ø, and thus buffers the inflection.





FIG. 4

is a partial cross sectional view of an alternative embodiment of an airfoil


100


that may be used with gas turbine engine


10


shown in FIG.


1


. Airfoil


100


is substantially similar to airfoil


40


shown in FIG.


3


and components in airfoil


100


that are identical to components of airfoil


40


are identified in

FIG. 3

using the same reference numerals used in FIG.


3


. Accordingly, airfoil


100


includes leading edge


48


, inflection


70


, and cooling chamber


58


. Airfoil


100


also includes a first sidewall


102


and a second sidewall


104


. Sidewalls


102


and


104


define cooling chamber


58


and are substantially similar to sidewalls


46


and


44


, shown in FIG.


3


.




A plurality of cooling openings


80


extend from cooling chamber


58


and airfoil outer surfaces


90


and


92


to connect cooling chamber


58


in flow communication with airfoil outer surfaces


90


and


92


. At least one cooling opening


110


extends between airfoil outer surface


90


and cooling chamber


58


within inflection


70


. More specifically, inflection cooling opening


110


has a centerline


112


and extends through sidewall


104


at an injection angle Ø. Injection angle Ø is formed by an intersection of centerline


112


and a line


114


that is tangent to airfoil outer surface


90


at a point where cooling opening


110


intersects airfoil outer surface


90


. In one embodiment, injection angle Ø is less than approximately 16 degrees. More specifically, because inflection cooling opening


110


extends through sidewall


104


, injection angle Ø is negative with respect to airfoil outer surface


90


. In an alternative embodiment, injection angle Ø is approximately equal to zero degrees.




During operation, because airfoil inflection


70


permits cooling fluid to be provided to airfoil outer surface


90


through inflection cooling opening


110


at a relatively shallow injection angle Ø, a reduction in injection jet separation is facilitated, thus enhancing film cooling effectiveness. Furthermore, because inflection


70


facilitates enhancing film cooling effectiveness, reduced amounts of cooling fluid for a set amount of heat transfer may be utilized. Alternatively, because inflection


70


facilitates enhancing film cooling effectiveness, a useful life of airfoil


100


may be facilitated to be extended.





FIG. 5

is a cross sectional view of an alternative embodiment of an airfoil


200


that may be used with a gas turbine engine, such as gas turbine engine


10


, shown in FIG.


1


. Airfoil


200


is substantially similar to airfoil


40


shown in FIG.


3


and components in airfoil


200


that are identical to components of airfoil


40


are identified in

FIG. 3

using the same reference numerals used in FIG.


3


. Accordingly, airfoil


200


includes leading edge


48


, inflection


70


, and cooling chamber


58


. Airfoil


200


also includes a first sidewall


202


and a second sidewall


204


. Sidewalls


202


and


204


define cooling chamber


58


and are substantially similar to sidewalls


44


and


46


, shown in

FIG. 3

, but sidewall


204


includes a plurality of inflections


208


. Inflections


208


extend longitudinally or radially between an airfoil root (not shown) and an airfoil tip (not shown), and are substantially similar to inflection


70


, but are formed within sidewall


204


.




At least one cooling opening


82


extends from cooling chamber


58


into inflection


70


. In an alternative embodiment, cooling opening


82


extends through either pressure side sidewall


202


or suction side sidewall


204


. More specifically, inflection cooling opening


82


has a centerline


84


, and extends through sidewall


202


at an injection angle Ø. Injection angle Ø is formed by an intersection of centerline


84


and tangential line


86


. In one embodiment, injection angle Ø is less than approximately 16 degrees.




A plurality of cooling openings


212


extend between cooling chamber


58


and airfoil outer surface


210


to connect cooling chamber


58


in flow communication with airfoil outer surface


210


. More specifically, each cooling opening


212


extends between airfoil outer surface


210


and cooling chamber


58


within a respective inflection


208


. More specifically, each cooling opening


212


has a centerline


214


, and extends through sidewall


204


at injection angle Ø. In one embodiment, each injection angle Ø is less than approximately 16 degrees. Each cooling opening


212


has a substantially circular diameter. Alternatively, cooling openings


212


are non-circular. In one embodiment, cooling openings


212


are cast with airfoil sidewall


204


and are not manufactured after casting of airfoil


200


. In another embodiment, cooling openings


212


are machined into airfoil


200


.




During operation, a velocity of combustion gases at and across airfoil leading edge


48


and airfoil pressure side sidewall


204


is relatively low in comparison to a velocity of the combustion gases across airfoil suction side sidewall


202


. As a result, low mach number velocity regions develop spaced axially from airfoil leading edge


48


along airfoil sidewall


204


, and higher mach number velocity regions develop downstream from leading edge


48


along airfoil sidewall


202


. Although film blowing ratios are typically higher in an airfoil low mach number velocity regions, because inflections


70


and


208


are formed within the airfoil low mach number velocity regions of airfoil


200


, cooling fluid is injected from cooling openings


82


and


212


, respectively, at a relatively shallow injection angle, and a reduction in film cooling separation is facilitated along airfoil suction sidewall


204


. In addition, because cooling fluid flow and injection angle are reduced along airfoil sidewall


202


, aerodynamic mixing losses are facilitated to be reduced.




The above-described airfoil includes at least one inflection and a cooling opening within the inflection. The inflection enables the inflection to extend from the cooling chamber with a relatively shallow injection angle to facilitate reducing aerodynamic mixing losses, and enhance film cooling effectiveness. As a result, enhanced film cooling facilitates extending a useful life of the airfoil in a cost-effective and reliable manner.




While the invention has been described in terms of various specific embodiments, those skilled in the art will recognize that the invention can be practiced with modification within the spirit and scope of the claims.



Claims
  • 1. A method for contouring an airfoil for a gas turbine engine to facilitate improving film cooling effectiveness of the airfoil, the airfoil including a leading edge, a trailing edge, a first sidewall, and a second sidewall, the first and second sidewalls connected chordwise at the leading and trailing edges to define a cavity, the first and second sidewalls extending radially between an airfoil root to an airfoil tip, said method comprising the steps of:forming an inflection in an outer surface of at least one of the airfoil first sidewall and the airfoil second sidewall, such that the inflection extends a distance radially between the airfoil root and the airfoil tip and such that the inflection is a distance downstream from the stagnation line of the leading edge of the airfoil, wherein the inflection is defined between a concave surface and a convex surface; and forming at least one opening immediately adjacent the inflection and through at least one of the convex surface and the concave surface immediately upstream of the inflection point of the inflection for receiving cooling fluid therethrough from the airfoil cavity to the airfoil outer surface wherein said opening through the airfoil sidewall is at an injection angle measured with respect to the airfoil outer surface that is less than about 16 degrees.
  • 2. A method in accordance with claim 1 wherein said step of extending each opening further comprises the step of extending each opening through the airfoil sidewall at an injection angle to reduce cooling flow to at least one of the airfoil first sidewall and the airfoil second sidewall.
  • 3. A method in accordance with claim 1 wherein said step of forming an inflection in an outer surface further comprises the step of forming a plurality of inflections in the airfoil outer surface.
  • 4. A method in accordance with claim 3 wherein the airfoil first side wall is substantially concave, and the airfoil second sidewall is substantially convex, said step of forming a plurality of inflections further comprises the steps of:forming at least one inflection in close proximity to the airfoil leading edge with, and forming at least one inflection within the airfoil second sidewall.
  • 5. An airfoil for a gas turbine engine, said airfoil comprising:a leading edge; a trailing edge; a first sidewall extending in radial span between an airfoil root and an airfoil tip, said first sidewall comprising an outer surface; a second sidewall connected to said first sidewall at said leading edge and said trailing edge, said second sidewall comprising an outer surface, and extending in radial span between the airfoil root and the airfoil tip, at least one of said first sidewall and said second side wall further comprising an inflection defined between a concave surface and a convex surface such that the inflection is a distance downstream of the stagnation line of the leading edge of the airfoil, at least one of said first sidewall and said second sidewall comprising at least one cooling opening extending therethrough immediately adjacent and upstream of the inflection point of said inflection, said at least one cooling opening configured to receive cooling fluid therethrough, wherein said at least one cooling opening is configured at an injection angle measured with respect to the airfoil outer surface that is less than about 16 degrees.
  • 6. An airfoil in accordance with claim 5 wherein each said cooling opening configured to reduce cooling flow to at least one of said airfoil first sidewall and said airfoil second sidewall.
  • 7. An airfoil in accordance with claim 5 wherein said airfoil first sidewall comprises a plurality of inflections, at least one of said inflections in close proximity to said airfoil leading edge.
  • 8. An airfoil in accordance with claim 7 wherein said airfoil first sidewall is substantially concave, said airfoil second sidewall is substantially convex.
  • 9. A gas turbine engine comprising a plurality of airfoils, each said airfoil comprising a leading edge, a trailing edge, a first sidewall comprising an outer surface, and a second sidewall comprising an outer surface, said airfoil first and second sidewalls connected chordwise at said leading and trailing edges, said first and second sidewalls extending radially from an airfoil root to an airfoil tip, at least one of said first sidewall and said second sidewall further comprising an inflection defined between a convex surface and a concave surface, wherein said inflection is a distance downstream of the stagnation line of the leading edge of the airfoil, at least one of said airfoil first and second sidewalls further comprises an opening extending therethrough immediately adjacent and upstream of the inflection point of said inflection, wherein said at least one cooling opening is configured at an injection angle measured with respect to the airfoil outer surface that is less than about 16 degrees.
  • 10. A gas turbine engine in accordance with claim 9 wherein each said airfoil first sidewall is substantially concave, each said airfoil second sidewall is substantially convex.
  • 11. A gas turbine engine in accordance with claim 10 wherein said airfoil first and second sidewalls define a cavity, said sidewall opening extending from said airfoil cavity to said airfoil outer surface.
  • 12. A gas turbine engine in accordance with claim 11 wherein each said airfoil sidewall opening configured to reduce cooling flow from said airfoil cavity to at least one of said airfoil first and second sidewalls.
  • 13. A gas turbine engine in accordance with claim 11 wherein at least one of said airfoil first and second sidewalls further comprises a plurality of inflections, at least one of said inflections in close proximity to said airfoil leading edge.
STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH & DEVELOPMENT

This invention was made with Government support under Contract No. F33615-92-C-2204 and Contract No. F33615-92-C-2278 awarded by the U.S. Air Force. The Government has certain rights in this invention.

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Webster's Third New International Dictionary, Unabridged, Copyright 1993 Merriam-Webster, Incorporated.