CA2709940A1 - Turbine airfoil casting method - Google Patents
Turbine airfoil casting method Download PDFInfo
- Publication number
- CA2709940A1 CA2709940A1 CA2709940A CA2709940A CA2709940A1 CA 2709940 A1 CA2709940 A1 CA 2709940A1 CA 2709940 A CA2709940 A CA 2709940A CA 2709940 A CA2709940 A CA 2709940A CA 2709940 A1 CA2709940 A1 CA 2709940A1
- Authority
- CA
- Canada
- Prior art keywords
- core
- airfoil
- core support
- alloy
- outer shell
- 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.)
- Abandoned
Links
Classifications
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/12—Blades
- F01D5/14—Form or construction
- F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
- F01D5/187—Convection cooling
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B22—CASTING; POWDER METALLURGY
- B22C—FOUNDRY MOULDING
- B22C9/00—Moulds or cores; Moulding processes
- B22C9/02—Sand moulds or like moulds for shaped castings
- B22C9/04—Use of lost patterns
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B22—CASTING; POWDER METALLURGY
- B22C—FOUNDRY MOULDING
- B22C9/00—Moulds or cores; Moulding processes
- B22C9/10—Cores; Manufacture or installation of cores
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B23—MACHINE TOOLS; METAL-WORKING NOT OTHERWISE PROVIDED FOR
- B23P—METAL-WORKING NOT OTHERWISE PROVIDED FOR; COMBINED OPERATIONS; UNIVERSAL MACHINE TOOLS
- B23P6/00—Restoring or reconditioning objects
- B23P6/04—Repairing fractures or cracked metal parts or products, e.g. castings
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2230/00—Manufacture
- F05D2230/20—Manufacture essentially without removing material
- F05D2230/21—Manufacture essentially without removing material by casting
- F05D2230/211—Manufacture essentially without removing material by casting by precision casting, e.g. microfusing or investment casting
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/202—Heat transfer, e.g. cooling by film cooling
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/221—Improvement of heat transfer
- F05D2260/2212—Improvement of heat transfer by creating turbulence
Abstract
A method for making a turbine airfoil includes: (a) providing a mold having.
(i) a core, (ii) an outer shell surrounding the core such that the core and the outer shell cooperatively define a cavity in the shape of an airfoil having at least one outer wall; and (Ui ) a core support extending from the core to the outer shell through a portion of the cavity that defines the at least one sidewalk (b) introducing molten metal alloy into the cavity and surrounding the core support; (c) solidifying the alloy to form an airfoil casting having at least one outer wall which has at least one core support opening passing therethrough; (d) removing the moid so as to expose the airfoil; and (e) sealing the at least one core support opening in the airfoil with a metal alloy mctallurgically bonded to the at least one outer wall.
(i) a core, (ii) an outer shell surrounding the core such that the core and the outer shell cooperatively define a cavity in the shape of an airfoil having at least one outer wall; and (Ui ) a core support extending from the core to the outer shell through a portion of the cavity that defines the at least one sidewalk (b) introducing molten metal alloy into the cavity and surrounding the core support; (c) solidifying the alloy to form an airfoil casting having at least one outer wall which has at least one core support opening passing therethrough; (d) removing the moid so as to expose the airfoil; and (e) sealing the at least one core support opening in the airfoil with a metal alloy mctallurgically bonded to the at least one outer wall.
Description
TURBINE 'c.I.RF OT C. CASTINQ :I -HITHOD
BACKGROUND OF THE INIVEN-TTON
[00011 This invention relates generally to the manufacture of gas turbine engine core ponertts and more particularly to methods for casting hollow turbine airfoils.
[000 1] Cast turbine airfoils fi r advanced gas turbine e r4giÃres have internal features that, challenge the capability of current casting technologies. The castings require complex ceramic cores to form the internal features and these cores are frag i le during the casting process, The result is that casting yields of 50 percent to 70 percent are not uncommon, The >0 percent to 50 percent casting scrap factors into the cost of the useable castings.
[0003] The issue is compounded by exotic alloys such as single crystal materials that drive up the cost to cast a part and thus drive up the cost caused by scrapping hardware,.
Ifa mere 5 percent to l 0 percent casting yield improvement can be achieved, the impact to each gas turbine engine is in the millions of dollars per year, based on volue.
[0004] One basic casting limitation is that the ceramic core that forms the internal structure of the airfoil can only be secured by the lower (i.e. root) portion with the majority of the core "floating" within the casting wax form. The forces of the molten metal and thermally induced forces during the cooling and solidification cycle result in movement and/or breakage of the ceramic core (referred to as "core sfrift").
The motion can be such that the cast component no longer meets drawing requirements, for example by violating minimum casting wall thicknesses. If the core fractures during the process this will also cause the component to fail requirements.
BRIE SL"MMA.RY OF'TIHT: INVFNTJON
[0005] These and other shortcomings of the prior art are addressed by the present invention, which provides a method fir supporting an airfoil core during casting while maintaining the z etallurgical integrity of the finished component.
[006] According to an aspect of the invention, a z netl od for making a turbine airfoil includes: (a) providing a mold having:(i) a core: (4) an outer shell surrounding the core _Ir such. that the core and the outer shell. cooperatively define a cavity in the shape of an airfoil having at least one outer wall; and 66)a core support extending from the core to the outer shell through a portion of the cavity that defines the at least one s dewall; (b) introducing molten metal alloy into the cavity and surrounding the core support; (c) solidifying the alloy to form an airfoil casting having at least one outer wall which has at least one core support opening passing therethrough; (d) removing the :mold so as to expose the airfoil; and (e) sealing the at least one core support opening:in the airfoil with a metal alloy metallurgically bonded to the at least one outer wall.
BRIEF D-11 SC'RIPT:ION OF THE DRAWINGS
[0007] The invention may be best understood by reference to the following description taken .in conjunction with the accompanying drawing figures in which:
[00081 Figure I is a perspective view of an exemplary turbine blade constructed in accordance with an aspect of the present invention;
[0009] Figure 2 is a perspective view of a mold core used in casting the blade shown in Figure 1, with a core support carried therein;
[00.10) Figure 3 is another perspective view of the mold core of Figure 2;
[0011] Figure 4 is a partial cross-sectional view of an assembled z Ãold;
[00121.1 Figure 5 is a crosssectional vier of the mold of Figure 4 with a portion of a blade casting therein;
[0Ã113] Figure 6 is a perspective view of an as-cast turbine bade, which includes an opening left by a core support;
[0014] Figure 7 is another perspective view of the turbine blade of Fi9tire 6;
[0015] Figure 8 is a partial cross-sectional view of the turbine blade taken along lines 8-8 of Figure 7;
[0016] Figure 9 is a cross--sectional view taken along lines 9--9 of Figure 1;
and [00.17] Figure 10 is a schematic view of an apparatus for closing the core support opening in the turbine blade, DETAILS ED DESCRIPTION IPTION O II1T_ INVENTI N
[0018] Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views, l;iggure .1 illustrates an exemplary, turbine blade 10. The turbine blade 10 includes a conventional dovetail 12, which may have any suitable torram including tangs that engage complementary tangs of a dovetail slot in a.
rotor disk (not shown; lfor radially retaining the blade 10 to the disk as it rotates during operation.. blade shank 14 extends radially, upwardly from the dovetail 1.2 and terminates in a platform 16 that projects laterally outwardly from and surrounds the shank 14. A hollow airfoil 18 extends radially outwardly from the platform 16.
The airfoil 18 has a concave pressure side outer wall 20 and a convex suction side outer wall 22 joined together at a leading edge 24 and at a trailing edge 26. The airfoil 18 may take any conh~..;uration suitable -for extracting energy from the hot as stream and causing rotation of the rotor disk. The blade 1.0 is preferably formed as a one-piece castino of a suitable "superalloy " of a knobs type, such as a nickel-based superalloy (c,g. Rene 80, Rene 142, ReneN 1, Rene N5) which has acceptable strength at the elevated temperatures of operation in a gas turbine en;ine. The airfoil 18 has a root 25 and a tip 27, and incorporates a member of trailing, ed4ge bleed holes 28.
[0019] The interior of the turbine blade 10 ismostly hollow and includesa nuniberof internal cooling features of a known type, such as walls defining serpentine passages, ribs, turbulence promoters ("turbulators"), etc. While the turbine blade 10 is a high pressure turbine blade, the principles of the present invention are applicable to any type of airfoil having a. holIoN interior.
[0020] Components such as the turbine blade 10 are raganti a.ctured using a known investment casting process. The method includes shaping the turbine blade in wax by enveloping a. conventional alumina or silica based ceramic core which defines internal coolant passages. The wax assembly then undergoes a series of dips in liquid ceramic solution. The part is allowed to dry after- each dip, forming a hard outer shell, typically a conventional zirconia based ceramic shell. After all dips are complete, and the wax assembly is encased by several layers of hardened ceramic shell, the asseirrbly is placed in a furnace where the wax in the shell is n relted out.
[00211 After wax rernova.l, the mold comprises the internal ceramic. core sur-r-ounded by the outer ceramic shell The cavity between the core and the outer shell defines the component and its interior features. The mold is again placed in the furnace, and liquid r aetal is poured into an opening at the top of the mold. The molten metal enters the space between the ceramic core and the ceramic shell, previously filled by the wax, After the metal is allowed to cool and solidify, the external shell is broken and removed, exposing the casting which has taken the shape of the cavity created by removal of the wax, and which encases the internal ceramic core. The castin4~ is then placed irr a leeching tank, where the core is dissolved, The component now has the shape of the wax form, and an internal cavity which was previously filled by the internal ceramic core.
[002. The relative thermal grog the the ceramic outer shell and the ceramic core material are different so that after the metal has been poured and is allowed to cool, the relative shrinking of the `hell and core components are different. This can cause varying wall thicknesses at areas of the metal nozzle part where one side of the wall is defined by the external shell, and the other side of the wall is engaged by the internal core.
Furthermore, the core is typically allowed to "float" and may thus shift it, position relative to the outer shell during the casting process. This can cause the walls of a.
component such as an airfoil to be less than a required minimum thickness.
[00.2 3 To avoid core shift, the turbine blade 10 is cast by a modification of the above process, which incorporates one or more core supports. Figures 2 and ;> are pre-casting views of a. core 3)0 with a core support 32 captured therein.. surrounding outer shell 34 comprises first and second sidewalls 34A and 34, as shown in Figure 4. Figure l also shows the core support 32 passing sequentially through the first side-"N--all 34A, a first portion J6 of wax fill, the core 30, a second portion of wax fill 38, and the.
second sidewall 34.H .
[0024] In the illustrated example, the core support 32 takes the form of a circular cross-section rod, but other cross-sectional. shapes may be used to suit a particular application.
[0025 The core support 32 is constructed from a suitable material having a meltin point higher than the alloy used for the turbine blade 10, which may be a known nickel-or cobalt-based "superalloy E: xa_mples of suitable core support materials include fused quartz, or a à erantic such as 4'ttri a, (Y 20,a) or samarium chide f San_C as used to mate the core 30.
[0026] ':1'he core support 32 remains in place during the casting process and resists motion of the core 30 during pouring and solidification. While any number of core supports 32 may be used and placed at any desired location, it is beneficial to support the core 30 in an area, denoted "A" in Figure 21, which defines the airfoil 1.8, This area of the core 30 is normally unsupported portion of the core ;its and is a stub tantial distance from the part of the core 30 which defines the blade shank 14. Support of the core helps maintain the core-to-outer shell spacing "S which directly affects the outer wall thickness of the finished turbine blade .10.
[0027] Figure 5 is a post-tasting partial cross-section which shows the dare support Q passing sequentially through the first sidewall 34A of the outer shell 34, the pressure, side outer wall 20 of the turbine blade 10, the core 30, the suction side outer wall 22. of the turbine blade 10, and the second sidewall 34B of the outer shell 34.
[0028] Figures 6-8 illustrate the turbine blade 10 after casting and removal of the outer shell 34, core. 30, and core support 32. The turbine blade 10 includes Core support openings 40 and 42 in the pressure and suction side outer walls 20 and 22;
respectively.
The core support openings 40 and 42 must be sealed before the turbine blade 10 is usable. Although it is possible to seal them using brazing techniques, this is not a metallurgical bond and does not have the same properties as the basic turbine blade i 0, which has a directionally-solidified or single-crystal microstructure imparting enhanced high-temperature strength and creep resistance.
[0029] An. example of a suitable apparatus, for scaling the core support openings 40 and _Sr 42 is disclosed in U. S. Patent N'u 5,622,638 to Schell et al, assigned to the assignee of this invention, and is schematically illustrated in :Figure .10. The apparatus includes a laser 44, an enclosed beam delivery conduit 46, laser focusing optics 48, a part positioning system 50, a vision system 52 for part location and laser path control, an optional preheat box (not shown), and a powder feed system 54 with a powder tube 56.
The working and coordination of the individual parts of the apparatus are controlled through a computerized system. controller 58. In a. conventional manner, the powder enters the laser beam in close proximity to the blade 10 as it is manipulated to cause melting and weld build-up.
[0030] The coree, support openings 40 and 42 may be sealed by using this apparatus to deposit molten alloy powder in one or more passes. Alternatively, powder can be deposited and then heated to melt and fuse it to the airfoil 18. In either case, the purer alloy composition is substantially the same as that of the basic turbine blade 1Ø This process, sometimes referred to as "reverse machining", produces a plug or patch that is metallurgically bonded to the core support openings 40 and 42, effectively forming an integral structure with. the turbine airfoil 10. With proper control of the Process parameters, this process can produce the same microstructure in the plug or patch (e. g.
directionally solidified or single crystal) as that of the turbine blade 10.
The finished turbine blade 1.0 is shown in Figures 1 and 9. This process will result in substantially higher casting yields, because, of the prevention of core shift, while maintaining the desired high-temperature, properties of the turbine blade 10.
[00311] The foregoing has described a method for making as turbine engine airfoils.
While specific embodiments of the present invention have been described, it will be apparent to those skilled in the art that various modifications thereto can be r :fade without departing from the spirit and scope of the invention. Accordingly, the foregoiffly description of the preferred embodiment of the invention and the best mode for practicing the invention are provided for the purpose of ill only and not for the purpose of limitation.
BACKGROUND OF THE INIVEN-TTON
[00011 This invention relates generally to the manufacture of gas turbine engine core ponertts and more particularly to methods for casting hollow turbine airfoils.
[000 1] Cast turbine airfoils fi r advanced gas turbine e r4giÃres have internal features that, challenge the capability of current casting technologies. The castings require complex ceramic cores to form the internal features and these cores are frag i le during the casting process, The result is that casting yields of 50 percent to 70 percent are not uncommon, The >0 percent to 50 percent casting scrap factors into the cost of the useable castings.
[0003] The issue is compounded by exotic alloys such as single crystal materials that drive up the cost to cast a part and thus drive up the cost caused by scrapping hardware,.
Ifa mere 5 percent to l 0 percent casting yield improvement can be achieved, the impact to each gas turbine engine is in the millions of dollars per year, based on volue.
[0004] One basic casting limitation is that the ceramic core that forms the internal structure of the airfoil can only be secured by the lower (i.e. root) portion with the majority of the core "floating" within the casting wax form. The forces of the molten metal and thermally induced forces during the cooling and solidification cycle result in movement and/or breakage of the ceramic core (referred to as "core sfrift").
The motion can be such that the cast component no longer meets drawing requirements, for example by violating minimum casting wall thicknesses. If the core fractures during the process this will also cause the component to fail requirements.
BRIE SL"MMA.RY OF'TIHT: INVFNTJON
[0005] These and other shortcomings of the prior art are addressed by the present invention, which provides a method fir supporting an airfoil core during casting while maintaining the z etallurgical integrity of the finished component.
[006] According to an aspect of the invention, a z netl od for making a turbine airfoil includes: (a) providing a mold having:(i) a core: (4) an outer shell surrounding the core _Ir such. that the core and the outer shell. cooperatively define a cavity in the shape of an airfoil having at least one outer wall; and 66)a core support extending from the core to the outer shell through a portion of the cavity that defines the at least one s dewall; (b) introducing molten metal alloy into the cavity and surrounding the core support; (c) solidifying the alloy to form an airfoil casting having at least one outer wall which has at least one core support opening passing therethrough; (d) removing the :mold so as to expose the airfoil; and (e) sealing the at least one core support opening:in the airfoil with a metal alloy metallurgically bonded to the at least one outer wall.
BRIEF D-11 SC'RIPT:ION OF THE DRAWINGS
[0007] The invention may be best understood by reference to the following description taken .in conjunction with the accompanying drawing figures in which:
[00081 Figure I is a perspective view of an exemplary turbine blade constructed in accordance with an aspect of the present invention;
[0009] Figure 2 is a perspective view of a mold core used in casting the blade shown in Figure 1, with a core support carried therein;
[00.10) Figure 3 is another perspective view of the mold core of Figure 2;
[0011] Figure 4 is a partial cross-sectional view of an assembled z Ãold;
[00121.1 Figure 5 is a crosssectional vier of the mold of Figure 4 with a portion of a blade casting therein;
[0Ã113] Figure 6 is a perspective view of an as-cast turbine bade, which includes an opening left by a core support;
[0014] Figure 7 is another perspective view of the turbine blade of Fi9tire 6;
[0015] Figure 8 is a partial cross-sectional view of the turbine blade taken along lines 8-8 of Figure 7;
[0016] Figure 9 is a cross--sectional view taken along lines 9--9 of Figure 1;
and [00.17] Figure 10 is a schematic view of an apparatus for closing the core support opening in the turbine blade, DETAILS ED DESCRIPTION IPTION O II1T_ INVENTI N
[0018] Referring to the drawings wherein identical reference numerals denote the same elements throughout the various views, l;iggure .1 illustrates an exemplary, turbine blade 10. The turbine blade 10 includes a conventional dovetail 12, which may have any suitable torram including tangs that engage complementary tangs of a dovetail slot in a.
rotor disk (not shown; lfor radially retaining the blade 10 to the disk as it rotates during operation.. blade shank 14 extends radially, upwardly from the dovetail 1.2 and terminates in a platform 16 that projects laterally outwardly from and surrounds the shank 14. A hollow airfoil 18 extends radially outwardly from the platform 16.
The airfoil 18 has a concave pressure side outer wall 20 and a convex suction side outer wall 22 joined together at a leading edge 24 and at a trailing edge 26. The airfoil 18 may take any conh~..;uration suitable -for extracting energy from the hot as stream and causing rotation of the rotor disk. The blade 1.0 is preferably formed as a one-piece castino of a suitable "superalloy " of a knobs type, such as a nickel-based superalloy (c,g. Rene 80, Rene 142, ReneN 1, Rene N5) which has acceptable strength at the elevated temperatures of operation in a gas turbine en;ine. The airfoil 18 has a root 25 and a tip 27, and incorporates a member of trailing, ed4ge bleed holes 28.
[0019] The interior of the turbine blade 10 ismostly hollow and includesa nuniberof internal cooling features of a known type, such as walls defining serpentine passages, ribs, turbulence promoters ("turbulators"), etc. While the turbine blade 10 is a high pressure turbine blade, the principles of the present invention are applicable to any type of airfoil having a. holIoN interior.
[0020] Components such as the turbine blade 10 are raganti a.ctured using a known investment casting process. The method includes shaping the turbine blade in wax by enveloping a. conventional alumina or silica based ceramic core which defines internal coolant passages. The wax assembly then undergoes a series of dips in liquid ceramic solution. The part is allowed to dry after- each dip, forming a hard outer shell, typically a conventional zirconia based ceramic shell. After all dips are complete, and the wax assembly is encased by several layers of hardened ceramic shell, the asseirrbly is placed in a furnace where the wax in the shell is n relted out.
[00211 After wax rernova.l, the mold comprises the internal ceramic. core sur-r-ounded by the outer ceramic shell The cavity between the core and the outer shell defines the component and its interior features. The mold is again placed in the furnace, and liquid r aetal is poured into an opening at the top of the mold. The molten metal enters the space between the ceramic core and the ceramic shell, previously filled by the wax, After the metal is allowed to cool and solidify, the external shell is broken and removed, exposing the casting which has taken the shape of the cavity created by removal of the wax, and which encases the internal ceramic core. The castin4~ is then placed irr a leeching tank, where the core is dissolved, The component now has the shape of the wax form, and an internal cavity which was previously filled by the internal ceramic core.
[002. The relative thermal grog the the ceramic outer shell and the ceramic core material are different so that after the metal has been poured and is allowed to cool, the relative shrinking of the `hell and core components are different. This can cause varying wall thicknesses at areas of the metal nozzle part where one side of the wall is defined by the external shell, and the other side of the wall is engaged by the internal core.
Furthermore, the core is typically allowed to "float" and may thus shift it, position relative to the outer shell during the casting process. This can cause the walls of a.
component such as an airfoil to be less than a required minimum thickness.
[00.2 3 To avoid core shift, the turbine blade 10 is cast by a modification of the above process, which incorporates one or more core supports. Figures 2 and ;> are pre-casting views of a. core 3)0 with a core support 32 captured therein.. surrounding outer shell 34 comprises first and second sidewalls 34A and 34, as shown in Figure 4. Figure l also shows the core support 32 passing sequentially through the first side-"N--all 34A, a first portion J6 of wax fill, the core 30, a second portion of wax fill 38, and the.
second sidewall 34.H .
[0024] In the illustrated example, the core support 32 takes the form of a circular cross-section rod, but other cross-sectional. shapes may be used to suit a particular application.
[0025 The core support 32 is constructed from a suitable material having a meltin point higher than the alloy used for the turbine blade 10, which may be a known nickel-or cobalt-based "superalloy E: xa_mples of suitable core support materials include fused quartz, or a à erantic such as 4'ttri a, (Y 20,a) or samarium chide f San_C as used to mate the core 30.
[0026] ':1'he core support 32 remains in place during the casting process and resists motion of the core 30 during pouring and solidification. While any number of core supports 32 may be used and placed at any desired location, it is beneficial to support the core 30 in an area, denoted "A" in Figure 21, which defines the airfoil 1.8, This area of the core 30 is normally unsupported portion of the core ;its and is a stub tantial distance from the part of the core 30 which defines the blade shank 14. Support of the core helps maintain the core-to-outer shell spacing "S which directly affects the outer wall thickness of the finished turbine blade .10.
[0027] Figure 5 is a post-tasting partial cross-section which shows the dare support Q passing sequentially through the first sidewall 34A of the outer shell 34, the pressure, side outer wall 20 of the turbine blade 10, the core 30, the suction side outer wall 22. of the turbine blade 10, and the second sidewall 34B of the outer shell 34.
[0028] Figures 6-8 illustrate the turbine blade 10 after casting and removal of the outer shell 34, core. 30, and core support 32. The turbine blade 10 includes Core support openings 40 and 42 in the pressure and suction side outer walls 20 and 22;
respectively.
The core support openings 40 and 42 must be sealed before the turbine blade 10 is usable. Although it is possible to seal them using brazing techniques, this is not a metallurgical bond and does not have the same properties as the basic turbine blade i 0, which has a directionally-solidified or single-crystal microstructure imparting enhanced high-temperature strength and creep resistance.
[0029] An. example of a suitable apparatus, for scaling the core support openings 40 and _Sr 42 is disclosed in U. S. Patent N'u 5,622,638 to Schell et al, assigned to the assignee of this invention, and is schematically illustrated in :Figure .10. The apparatus includes a laser 44, an enclosed beam delivery conduit 46, laser focusing optics 48, a part positioning system 50, a vision system 52 for part location and laser path control, an optional preheat box (not shown), and a powder feed system 54 with a powder tube 56.
The working and coordination of the individual parts of the apparatus are controlled through a computerized system. controller 58. In a. conventional manner, the powder enters the laser beam in close proximity to the blade 10 as it is manipulated to cause melting and weld build-up.
[0030] The coree, support openings 40 and 42 may be sealed by using this apparatus to deposit molten alloy powder in one or more passes. Alternatively, powder can be deposited and then heated to melt and fuse it to the airfoil 18. In either case, the purer alloy composition is substantially the same as that of the basic turbine blade 1Ø This process, sometimes referred to as "reverse machining", produces a plug or patch that is metallurgically bonded to the core support openings 40 and 42, effectively forming an integral structure with. the turbine airfoil 10. With proper control of the Process parameters, this process can produce the same microstructure in the plug or patch (e. g.
directionally solidified or single crystal) as that of the turbine blade 10.
The finished turbine blade 1.0 is shown in Figures 1 and 9. This process will result in substantially higher casting yields, because, of the prevention of core shift, while maintaining the desired high-temperature, properties of the turbine blade 10.
[00311] The foregoing has described a method for making as turbine engine airfoils.
While specific embodiments of the present invention have been described, it will be apparent to those skilled in the art that various modifications thereto can be r :fade without departing from the spirit and scope of the invention. Accordingly, the foregoiffly description of the preferred embodiment of the invention and the best mode for practicing the invention are provided for the purpose of ill only and not for the purpose of limitation.
Claims (12)
1. A method for making a turbine airfoil, comprising:
(a) providing a mold having:
(i) a core;
(ii) an outer shell surrounding the core such that the core and the outer shell cooperatively define a cavity in the shape of an airfoil having at least one outer wall; and (iii) a core support extending from the core to the outer shell through a portion of the cavity that defines the at least one sidewall;
(b) introducing molten metal alloy into the cavity and surrounding the core support;
(c) solidifying the alloy to form an airfoil casting having at least one outer wall which has at least one core support opening passing therethrough;
(d) removing the mold so as to expose the airfoil; and (e) sealing the at least one core support opening in the airfoil with a metal alloy metallurgically bonded to the at least one outer wall.
(a) providing a mold having:
(i) a core;
(ii) an outer shell surrounding the core such that the core and the outer shell cooperatively define a cavity in the shape of an airfoil having at least one outer wall; and (iii) a core support extending from the core to the outer shell through a portion of the cavity that defines the at least one sidewall;
(b) introducing molten metal alloy into the cavity and surrounding the core support;
(c) solidifying the alloy to form an airfoil casting having at least one outer wall which has at least one core support opening passing therethrough;
(d) removing the mold so as to expose the airfoil; and (e) sealing the at least one core support opening in the airfoil with a metal alloy metallurgically bonded to the at least one outer wall.
2. The method of claim 1 wherein step (e) is carried out by:
(a) melting an alloy power using laser energy;
(b) introducing the melted powder into the at least one core support opening;
and (c) allowing the melted powder to cool and solidify in the at least one core support opening.
(a) melting an alloy power using laser energy;
(b) introducing the melted powder into the at least one core support opening;
and (c) allowing the melted powder to cool and solidify in the at least one core support opening.
3. The method of claim 2 wherein the alloy power has a composition substantially the same as an alloy composition of the airfoil.
4. The method of claim 1 wherein step (e) is carried out by:
(a) introducing an alloy powder into the at least one core support opening;
(b) using laser energy, melting the alloy powder and heating the outer wall;
and (c) allowing the melted powder to cool and solidify in the at least one core support opening.
(a) introducing an alloy powder into the at least one core support opening;
(b) using laser energy, melting the alloy powder and heating the outer wall;
and (c) allowing the melted powder to cool and solidify in the at least one core support opening.
5. The method of claim 4 wherein the alloy power has a composition substantially the same as an alloy composition of the airfoil.
6. The method of claim 1 wherein the airfoil has a directionally solidified or single crystal microstructure, and the solidified powder in the at least one core support opening has substantially the same microstructure as the airfoil.
7. The method of claim 1 wherein the core support is a cylindrical rod.
8. The method of claim 1 wherein the core support comprises fused quartz.
9. The method of claim 1 wherein the mold comprises:
(i) a core; and (ii) an outer shell surrounding the core, the outer shell having two spaced-apart sidewalls disposed on opposite sides of the core; and (iii) a core support extending from the one of the sidewalls of the outer shell, through the core, to the other sidewall of the outer shell.
(i) a core; and (ii) an outer shell surrounding the core, the outer shell having two spaced-apart sidewalls disposed on opposite sides of the core; and (iii) a core support extending from the one of the sidewalls of the outer shell, through the core, to the other sidewall of the outer shell.
10. The method of claim 9 wherein the airfoil includes spaced-apart pressure and suction side outer walls each having a core support opening therein, the core support openings being coaxial to each other.
11. The method of claim 1 wherein the mold defines an integral turbine blade including the airfoil and a shank, the airfoil and the shank respectively defining opposite ends of the turbine blade.
12. The method of claim 11 wherein the core support is positioned in the airfoil at an end of the turbine blade opposite from the shank.
Applications Claiming Priority (3)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| US11/968,022 US20090165988A1 (en) | 2007-12-31 | 2007-12-31 | Turbine airfoil casting method |
| US11/968,022 | 2007-12-31 | ||
| PCT/US2008/086569 WO2009085654A1 (en) | 2007-12-31 | 2008-12-12 | Turbine airfoil casting method |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CA2709940A1 true CA2709940A1 (en) | 2009-07-09 |
Family
ID=40342316
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CA2709940A Abandoned CA2709940A1 (en) | 2007-12-31 | 2008-12-12 | Turbine airfoil casting method |
Country Status (6)
| Country | Link |
|---|---|
| US (1) | US20090165988A1 (en) |
| JP (1) | JP2011509185A (en) |
| CA (1) | CA2709940A1 (en) |
| DE (1) | DE112008003545T5 (en) |
| GB (1) | GB2468083A (en) |
| WO (1) | WO2009085654A1 (en) |
Families Citing this family (17)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US8752609B2 (en) * | 2009-11-05 | 2014-06-17 | Dresser-Rand Company | One-piece manufacturing process |
| US9550230B2 (en) * | 2011-09-16 | 2017-01-24 | United Technologies Corporation | Mold for casting a workpiece that includes one or more casting pins |
| SG10201900946WA (en) | 2012-12-14 | 2019-03-28 | United Technologies Corp | Hybrid turbine blade for improved engine performance or architecture |
| SG10201610144XA (en) | 2012-12-14 | 2017-01-27 | United Technologies Corp | Multi-shot casting |
| US9057276B2 (en) | 2013-02-06 | 2015-06-16 | Siemens Aktiengesellschaft | Twisted gas turbine engine airfoil having a twisted rib |
| US9797258B2 (en) * | 2013-10-23 | 2017-10-24 | General Electric Company | Turbine bucket including cooling passage with turn |
| WO2015112583A1 (en) * | 2014-01-21 | 2015-07-30 | United Technologies Corporation | Method for forming single crystal components using additive manufacturing and re-melt |
| US9435211B2 (en) | 2014-05-09 | 2016-09-06 | United Technologies Corporation | Method for forming components using additive manufacturing and re-melt |
| FR3037972B1 (en) * | 2015-06-29 | 2017-07-21 | Snecma | PROCESS SIMPLIFYING THE CORE USED FOR THE MANUFACTURE OF A TURBOMACHINE BLADE |
| FR3037830B1 (en) * | 2015-06-29 | 2024-02-16 | Snecma | MOLDING ASSEMBLY FOR A TURBOMACHINE BLADE, INCLUDING A LARGE SECTION RELIEF PORTION |
| US10053989B2 (en) | 2015-12-21 | 2018-08-21 | General Electric Company | Cooling circuit for a multi-wall blade |
| US11813669B2 (en) | 2016-12-13 | 2023-11-14 | General Electric Company | Method for making an integrated core-shell structure |
| US20180161853A1 (en) * | 2016-12-13 | 2018-06-14 | General Electric Company | Integrated casting core-shell structure with floating tip plenum |
| US10807154B2 (en) * | 2016-12-13 | 2020-10-20 | General Electric Company | Integrated casting core-shell structure for making cast component with cooling holes in inaccessible locations |
| US20180161866A1 (en) | 2016-12-13 | 2018-06-14 | General Electric Company | Multi-piece integrated core-shell structure for making cast component |
| WO2019046036A1 (en) * | 2017-08-28 | 2019-03-07 | Siemens Aktiengesellschaft | Method for making a turbine airfoil |
| DE102021204782A1 (en) | 2021-05-11 | 2022-11-17 | Siemens Energy Global GmbH & Co. KG | Improved blade tip in new or repaired part and process |
Family Cites Families (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US3659645A (en) * | 1965-08-09 | 1972-05-02 | Trw Inc | Means for supporting core in open ended shell mold |
| US5810552A (en) * | 1992-02-18 | 1998-09-22 | Allison Engine Company, Inc. | Single-cast, high-temperature, thin wall structures having a high thermal conductivity member connecting the walls and methods of making the same |
| US5584663A (en) | 1994-08-15 | 1996-12-17 | General Electric Company | Environmentally-resistant turbine blade tip |
| DE19905887C1 (en) * | 1999-02-11 | 2000-08-24 | Abb Alstom Power Ch Ag | Hollow cast component |
| US6454156B1 (en) * | 2000-06-23 | 2002-09-24 | Siemens Westinghouse Power Corporation | Method for closing core printout holes in superalloy gas turbine blades |
| US6495793B2 (en) * | 2001-04-12 | 2002-12-17 | General Electric Company | Laser repair method for nickel base superalloys with high gamma prime content |
| EP1396299A1 (en) * | 2002-09-06 | 2004-03-10 | Siemens Aktiengesellschaft | Process for filling up a hole in a component |
| EP1561536A1 (en) * | 2004-02-03 | 2005-08-10 | Siemens Aktiengesellschaft | Process of brazing repairing of a part having a base material with oriented microstructure |
| US7322396B2 (en) * | 2005-10-14 | 2008-01-29 | General Electric Company | Weld closure of through-holes in a nickel-base superalloy hollow airfoil |
-
2007
- 2007-12-31 US US11/968,022 patent/US20090165988A1/en not_active Abandoned
-
2008
- 2008-12-12 WO PCT/US2008/086569 patent/WO2009085654A1/en active Application Filing
- 2008-12-12 JP JP2010540763A patent/JP2011509185A/en active Pending
- 2008-12-12 GB GB1010159A patent/GB2468083A/en not_active Withdrawn
- 2008-12-12 DE DE112008003545T patent/DE112008003545T5/en not_active Withdrawn
- 2008-12-12 CA CA2709940A patent/CA2709940A1/en not_active Abandoned
Also Published As
| Publication number | Publication date |
|---|---|
| DE112008003545T5 (en) | 2010-12-09 |
| JP2011509185A (en) | 2011-03-24 |
| GB201010159D0 (en) | 2010-07-28 |
| GB2468083A (en) | 2010-08-25 |
| WO2009085654A1 (en) | 2009-07-09 |
| US20090165988A1 (en) | 2009-07-02 |
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