US20180015536A1 - Forming cooling passages in combustion turbine superalloy castings - Google Patents
Forming cooling passages in combustion turbine superalloy castings Download PDFInfo
- Publication number
- US20180015536A1 US20180015536A1 US15/548,267 US201615548267A US2018015536A1 US 20180015536 A1 US20180015536 A1 US 20180015536A1 US 201615548267 A US201615548267 A US 201615548267A US 2018015536 A1 US2018015536 A1 US 2018015536A1
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- ceramic
- component
- insert
- cooling passage
- shell
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- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
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- F04—POSITIVE - DISPLACEMENT MACHINES FOR LIQUIDS; PUMPS FOR LIQUIDS OR ELASTIC FLUIDS
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- B32B2260/00—Layered product comprising an impregnated, embedded, or bonded layer wherein the layer comprises an impregnation, embedding, or binder material
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
- Y02T50/00—Aeronautics or air transport
- Y02T50/60—Efficient propulsion technologies, e.g. for aircraft
Definitions
- the invention relates to methods for forming cooling passages in superalloy substrates of combustion turbine components, such as engine blades, vanes, or transitions. More particularly, the invention relates to formation of partially completed cooling passages in such superalloy components during investment casting, by use of ceramic shell inserts, with projecting ceramic posts, then completing the passages in the castings.
- Known turbine engines including gas/combustion turbine engines and steam turbine engines, incorporate shaft-mounted turbine blades circumferentially circumscribed by a turbine casing or housing.
- hot combustion gasses flow in a combustion path that initiates within a combustor and are directed through a generally tubular transition into a turbine section.
- a forward or Row 1 vane directs the combustion gasses past successive alternating rows of turbine blades and vanes. Hot combustion gas striking the turbine blades cause blade rotation, thereby converting thermal energy within the hot gasses to mechanical work, which is available for powering rotating machinery, such as an electrical generator.
- Engine internal components within the hot combustion gas path are exposed to combustion temperatures approximately well over 1000 degrees Celsius (1832 degrees Fahrenheit).
- the engine internal components within the combustion path such as for example combustion section transitions, vanes and blades are often constructed of high temperature resistant superalloys. Blades, vanes, and transitions often include cooling passages terminating in cooling holes on component outer surface, for passage of coolant fluid into the combustion path.
- Turbine engine internal components often incorporate a thermal barrier coat or coating (“TBC”) of metal-ceramic material that is applied directly to the external surface of the component substrate surface or over an intermediate metallic bond coat (“BC”) that was previously applied to the substrate surface.
- TBC thermal barrier coat or coating
- the TBC provides a thermal insulating layer over the component substrate, which reduces the substrate temperature. Combination of TBC application along with cooling passages in the component further lowers the substrate temperature.
- cooling passages are formed by removing superalloy material from the intended passage path within the component, with exemplary removal tools including mechanical cutting/drilling bits, or various ablation devices, such as high-pressure water jet, percussion laser pulsation, and electric discharge machining (“EDM”).
- exemplary removal tools including mechanical cutting/drilling bits, or various ablation devices, such as high-pressure water jet, percussion laser pulsation, and electric discharge machining (“EDM”).
- EDM electric discharge machining
- Cut cooling passage path, profile and size are limited by the physical capabilities of the cutting instrument. For example, drilled passages are linear and have cross sectional symmetry to match the drill bit.
- Ablated passages are limited by the size of the ablation instrument and ability to maneuver the instrument along a cutting path.
- Investment cast turbine engine components are fabricated by creating a hardened wax pattern, in a wax injection mold, which replicates the profile of the finished superalloy component.
- the wax pattern is enveloped in ceramic slurry, which is subsequently hardened by firing, into a ceramic shell casing.
- the internal cavity is filled with molten superalloy material.
- wax patterns for investment cast, superalloy components for combustion turbine engines are injected into hard tool wax molds, and removed from the tools with precise and smooth surfaces.
- the wax patterns are then dipped in various ceramic slurry mixtures and processed to form the ceramic outer shell, which is subsequently sintered to form a vestibule in which molten metal is poured.
- the outer ceramic shell is removed by mechanical and/or chemical methods and the metal part is then prepared for further processing. Further processing of the metal part includes ceramic core removal, finish machining, drilling of cooling holes, and application of a thermal barrier coating (“TBC”).
- TBC thermal barrier coating
- Current state of the art processes often require the investment cast surface be lightly grit blasted to prepare the surface for bond coat application.
- a bond coat typically a metallic Cramoium, Aluminum, Yitria (“MCrAlY”) coating is applied to the substrate via a spray deposition technique, such as High Velocity Oxy Fuel (“HVOF”) or Low Pressure Plasma Spray (“LPPS”).
- HVOF High Velocity Oxy Fuel
- LPPS Low Pressure Plasma Spray
- a ceramic thermal barrier such as YSZ (Yttria Stabilized Zirconia) is applied to the surface of the MCrAlY via atmospheric or air plasma spray (“APS”) to complete the coating system.
- APS atmospheric or air plasma spray
- a two layer ceramic coating is applied via APS for low thermal conductivity.
- the investment casting wax pattern does not have sufficient, reliable, structural integrity to form cooling passages directly therein.
- cooling passages are formed in the mold that forms the wax pattern, there is more than insignificant chance that the cooling passage profile in the wax pattern will deform, or that the wax pattern passage will not fill completely with ceramic slurry; in either case the resultant passage in the metal casting does not confirm to design specification.
- RMC refractory metal core
- cooling passages are formed in the blade, vane, transition, or other superalloy component prior to application of the TBC layer, the passages will become obstructed by the TBC material as the latter is applied to the component surface. Obstruction can be mitigated by temporarily masking the cooling passages on the component surface prior to the TBC application, which adds additional, costly, steps to the manufacturing processes. In the alternative, excess TBC material obstructions within cooling passages can be removed subsequently by the aforementioned cutting processes. Post TBC-application cooling passage obstruction removal increases risk of TBC layer damage and/or delamination along the margins of cooling passages on the component surface.
- cooling passages are formed after application of a TBC layer to the component substrate. In one known post TBC-coating cooling passage formation process, a pulsed laser ablates TBC material from the component at the intended cooling passage entry point, and then ablates the superalloy material to form the passage.
- a ceramic insert which incorporates an engineered surface with one or more projecting posts, conforming to respective partially completed cooling passage(s), is incorporated into the casting mold.
- the ceramic insert is a partially sintered, ceramic material (similar to typical core material used in investment casting processes) that is positioned onto and/or embedded within the wax pattern.
- the ceramic shell insert is placed into the wax injection tool and embedded within the wax pattern when it is injected.
- Exemplary embodiments of the methods described herein form cooling passages in combustion turbine engine components, such as blades, vanes, or transitions, during investment casting, through use of ceramic shell inserts within the casting mold.
- Ceramic posts formed in the ceramic shell insert have profiles conforming to profiles of partially completed cooling passages.
- the shell insert surface, including its posts, is embedded in a hardened wax pattern, whose profile conforms to the component profile.
- the wax pattern and its embedded shell insert are enveloped in ceramic slurry that is hardened to form an outer ceramic shell. After removal of the hardened wax, the cavity within the outer ceramic shell is filled with melted superalloy material, and thereafter hardened to form a superalloy component casting.
- the ceramic outer shell, shell insert and any cores are removed from the casting, which now has cast-in-place, partial cooling passages formed therein.
- the cooling passages are completed by removing remaining superalloy material in the cooling passage path.
- Passage path profile is selectively varied easily by varying profile of the ceramic insert posts, which are affixed to the ceramic insert, and which do not need additional alignment effort within the casting molds, compared to standard, known RMC inserts.
- the partially completed cooling path passages are easily cleared of any subsequently applied TBC layer obstructions, prior to completion of the passages.
- the cooling passage is completed by removing the rest of the superalloy material from the casting to match a corresponding, completed cooling passage profile in the corresponding component.
- Exemplary embodiments of the invention feature a method for forming a cooling passage in an investment cast, superalloy component for a combustion turbine engine, by providing a wax injection mold defining a mold cavity, whose mold cavity surface conforms to a corresponding surface profile of a component for a combustion turbine engine.
- a ceramic shell insert is provided, having an insert surface profile and at least one ceramic post projecting from the inner surface, which in combination conform to a corresponding surface profile of a partially completed cooling passage in the engine component.
- the ceramic shell insert is inserted into the wax injection mold.
- the shell insert surface and each ceramic post forms part of the mold cavity surface, with each post projecting into the mold cavity.
- the mold cavity is filled with wax, enveloping each post therein.
- the wax is hardened, creating a wax pattern that embeds the ceramic shell and each post therein.
- the wax injection mold is removed.
- the hardened wax pattern, along with the insert surface and each ceramic post conforms to the component surface profile.
- the hardened wax pattern and embedded shell insert are enveloped in ceramic slurry.
- the ceramic slurry is fired, thereby hardening the slurry into an outer ceramic shell, which is joined to the ceramic shell insert, and eliminating the wax.
- the joined outer ceramic shell and ceramic shell insert define a shell internal cavity, which conforms to the engine component profile, including each partially completed cooling passage.
- the shell internal cavity is filled with molten superalloy material.
- the superalloy material is cooled, to form a component casting therein.
- the casting includes each partially completed cooling passage.
- the outer ceramic shell and ceramic insert, including each ceramic post, are removed from the casting, exposing each partially completed cooling passage.
- At least one cooling passage is completed by removing hardened, cast superalloy material from the casting to match a corresponding, completed cooling passage profile in the corresponding component.
- exemplary embodiments of the invention feature a method for forming a cooling passage in an investment cast, superalloy blade or vane component for a combustion turbine engine.
- the component has a component wall delimited by a wall outer surface, and a wall inner surface that defines a component cavity therein, with the cooling passage extending through the component wall between its respective outer and inner surfaces.
- a wax injection mold is provided, defining a mold cavity, whose mold cavity surface conforms to a corresponding profile of an outer surface of a hollow blade or vane component for a combustion turbine engine.
- a ceramic shell insert is provided, having an insert surface profile and at least one ceramic post projecting from the insert surface, which in combination conform to a corresponding surface profile of a partially completed cooling passage in the engine component wall.
- a ceramic core is provided, whose core outer surface conforms to a corresponding profile of a wall inner surface of the same component.
- the ceramic core is inserted into the mold cavity, and the ceramic shell insert is inserted into the mold, with the shell insert surface in opposed, spaced relationship with the ceramic core outer surface, forming a cavity void there between, which corresponds to a corresponding profile of the component wall respective outer and inner surfaces.
- the mold cavity void is filled with wax, enveloping each ceramic post therein.
- the wax is hardened, forming a wax pattern that embeds the ceramic core, and ceramic shell insert, including each ceramic post therein.
- the wax injection mold is removed, with the hardened wax pattern, along with the ceramic insert surface and each ceramic post conforming to the component wall respective outer and inner surfaces profiles.
- the hardened wax pattern and embedded shell insert are enveloped in ceramic slurry, which is fired, thereby hardening the slurry into an outer ceramic shell, which is joined to the ceramic shell insert and the ceramic core.
- the wax is eliminated, with the joined outer ceramic shell and ceramic insert defining a shell internal cavity, which conforms to the engine component wall outer surface profile, including each partially completed cooling passage.
- the ceramic core outer surface conforms to the engine component wall inner surface profile
- the ceramic shell cavity void conforms to the engine component wall.
- the ceramic shell internal cavity void is filled with molten superalloy material, which is then cooled to form a component casting therein.
- the casting includes each partially completed cooling passage in the component outer wall surface.
- the ceramic core, ceramic outer shell, and ceramic insert, including each ceramic post, are removed from the casting, exposing each partially completed cooling passage in the outer surface of the component wall. At least one cooling passage is completed through the component wall, by removing hardened, cast superalloy material that blocks remaining path of the cooling passage.
- FIG. 1 is a partial axial cross sectional view of a gas or combustion turbine engine incorporating one or more superalloy components, having cooling passages formed in accordance with exemplary method embodiments of the invention
- FIG. 2 is a detailed cross sectional elevational view of the turbine engine of FIG. 1 , showing Rows 1 turbine blade and Rows 1 and 2 vanes, having cooling passages formed in accordance with exemplary method embodiments of the invention;
- FIG. 3 is an elevational perspective view of an exemplary ceramic shell insert, having ceramic posts and engineered surface features (“ESFs”), which is used to form cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention;
- ESFs engineered surface features
- FIG. 4 is a perspective view of exemplary ceramic posts and ESFs of the ceramic shell insert of FIG. 3 ;
- FIG. 5 is a cross-sectional plan view of a wax injection mold for a turbine blade airfoil, incorporating the ceramic shell insert of FIG. 3 and a ceramic core, during wax injection into the mold cavity, which is used to form cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention;
- FIG. 6 is a cross-sectional plan view of a hardened wax pattern, with embedded ceramic shell insert and ceramic core, for a turbine blade airfoil, after removal from the wax injection mold of FIG. 5 , which is used to form cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention;
- FIG. 7 is a cross-sectional plan view of the ceramic shell insert and ceramic core after envelopment in an outer ceramic shell and removal of the hardened wax, which is used to form cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention
- FIG. 8 is a cross-sectional view of a ceramic shell insert and ceramic post, which is used to form cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention.
- FIG. 9 is a cross-sectional view of the ceramic shell insert and ceramic post of FIG. 8 , in opposed orientation with the ceramic core, after wax injection, in accordance with exemplary method embodiments of the invention.
- FIG. 10 is a cross-sectional view of the ceramic shell insert and ceramic post of FIG. 8 , in opposed orientation with the ceramic core, after envelopment in an outer ceramic shell and removal of the hardened wax, in accordance with exemplary method embodiments of the invention;
- FIG. 11 is a cross-sectional view of the ceramic shell insert and ceramic post of FIG. 8 , in opposed orientation with the ceramic core, after filling the mold with molten superalloy material and subsequent hardening of the material, in accordance with exemplary method embodiments of the invention;
- FIG. 12 is a cross-sectional elevation view of the superalloy component casting, after removal of the ceramic shell insert, ceramic core and ceramic outer shell, and a formed, partial depth cooling passage formed therein, in accordance with exemplary method embodiments of the invention
- FIG. 13 is a cross-sectional view of a ceramic shell insert and alternative embodiment converging frustro-conical profile ceramic post, in opposed orientation with a ceramic core, similar to FIG. 8 , which is used to form cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention;
- FIG. 14 is a cross-sectional view of a ceramic shell insert and alternative embodiment diverging profile ceramic post, in opposed orientation with a ceramic core, similar to FIG. 8 , which is used to form cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention;
- FIG. 15 is a cross-sectional view of a ceramic shell insert and alternative embodiment profile rectangular profile ceramic post, in opposed orientation with a ceramic core, similar to FIG. 8 , which is used to form cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention;
- FIG. 16 is a cross-sectional view of a ceramic shell insert and alternative embodiment trapezoidal profile ceramic post, in opposed orientation with a ceramic core, similar to FIG. 8 , which is used to form cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention;
- FIGS. 17 and 18 are cross-sectional views of ceramic shell inserts and alternative embodiment trapezoidal profile ceramic posts in opposed orientation with ceramic cores, similar to FIG. 8 , which are used to form film cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention;
- FIGS. 19 and 20 are cross-sectional views of ceramic shell inserts and alternative embodiment split-profile ceramic posts in opposed orientation with ceramic cores, similar to FIG. 8 , which are used to form split cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention;
- FIGS. 21 and 22 are cross-sectional views of ceramic shell inserts and alternative embodiment non-linear, asymmetrical profile ceramic posts in opposed orientation with ceramic cores, similar to FIG. 8 , which are used to non-linear and/or asymmetrical cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention;
- FIGS. 23-25 are elevational cross section views of a prior art method for forming cooling passages in a thermal barrier coated (“TBC”) superalloy component for a combustion turbine engine, leading to undesired TBC delamination from the component substrate and its bond coat (“BC”) layer around the passage margins; and
- TBC thermal barrier coated
- FIGS. 26 and 27 are elevational cross section views of an exemplary method for forming cooling passages in a thermal barrier coated (“TBC”) superalloy component for a combustion turbine engine, where a previously formed, partial cooling passage inhibits undesired TBC delamination from the component substrate and its bond coat (“BC”) layer around the passage margin, by shielding the TBC layer with an overhanging layer of superalloy material, in accordance with exemplary method embodiments of the invention.
- TBC thermal barrier coated
- Exemplary method embodiments of the invention form partially completed cooling passages directly in turbine engine components during investment casting of the superalloy material.
- the partial cooling passages are formed by use of ceramic shell inserts, having projecting posts, and optionally additional engineered surface features.
- the ceramic posts have sufficient strength to maintain structural integrity and alignment as they are being embedded in wax during the wax pattern formation of the manufacturing process.
- the composite hardened wax pattern and embedded ceramic shell, along with any other casting mold segments, such as central cores are in turn enveloped in ceramic slurry.
- the slurry is fired, hardening it into an outer ceramic shell. Any wax that was not previously eliminated during the outer ceramic shell firing is removed, leaving a mold cavity, for receipt of molten superalloy material.
- the ceramic posts of the ceramic shell insert project into the mold cavity.
- Molten superalloy material poured into the mold cavity hardens around the ceramic posts. After the casting cools and hardens, the ceramic mold material, including the outer shell, shell insert and the projecting ceramic posts are removed; leaving partially formed or completed cooling passages in the component substrate.
- a TBC layer is applied over desired portions of the component substrate, including the partially completed cooling passages. Any TBC material obstructing the partial cooling passages is removed, allowing access to the terminus of the partial cooling passage. Thereafter the cooling passage path is completed by removing component superalloy material along the intended passage path.
- turbine engines such as the gas or combustion turbine engine 80 include a multi-stage compressor section 82 , a combustion section 84 , a multi-stage turbine section 86 and an exhaust system 88 .
- Atmospheric pressure intake air is drawn into the compressor section 82 generally in the direction of the flow arrows F along the axial length of the turbine engine 80 .
- the intake air is progressively pressurized in the compressor section 82 by rows rotating compressor blades and directed by mating compressor vanes to the combustion section 84 , where it is mixed with fuel and ignited.
- the ignited fuel/air mixture now under greater pressure and velocity than the original intake air, is directed through a transition 85 to the sequential blade rows R 1 , R 2 , etc., in the turbine section 86 .
- the engine's rotor and shaft 90 has a plurality of rows of airfoil cross sectional shaped turbine blades 92 terminating in distal blade tips 94 in the compressor 82 and turbine 86 sections.
- each turbine blade 92 has a concave profile high-pressure side 96 and a convex low-pressure side 98 .
- Cooling passages 99 that are formed in the blade 92 facilitate passage of cooling fluid along the blade surface. The high velocity and pressure combustion gas, flowing in the combustion flow direction F imparts rotational motion on the blades 92 , spinning the rotor 90 .
- combustion gasses are constrained radially distal the rotor 90 by turbine casing 100 and proximal the rotor 90 by air seals 102 comprising abradable surfaces.
- respective upstream vanes 104 and downstream vanes 106 respectively direct upstream combustion gas generally parallel to the incident angle of the leading edge of turbine blade 92 and redirect downstream combustion gas exiting the trailing edge of the blade 92 for a desired entry angle into downstream Row 2 turbine blades (not shown).
- Cooling passages 105 that are formed in the vanes 104 , 106 facilitate passage of cooling fluid along the vane surface. It is noted that the cooling passages 99 and 105 shown in FIG. 2 are merely schematic representations, are enlarged for visual clarity, and are not drawn to scale.
- a typical turbine blade 92 or vane 104 , 106 has many more cooling passages distributed about the respective airfoil bodies of much smaller diameter relative to the respective blade or vane total surface area that is exposed to the engine combustion gas.
- turbine component surfaces that are exposed to combustion gasses are often constructed with a TBC layer for insulation of their underlying substrates.
- Typical TBC coated surfaces include the turbine blades 92 , the vanes 104 and 106 , ring segments 110 , abradable surfaces 120 and related carrier surfaces of turbine vanes, and combustion section transitions 85 .
- the TBC layer for blade 92 , vanes 104 and 106 , ring segments 110 , and transition 85 exposed surfaces are often applied by thermal sprayed or vapor deposition or solution/suspension plasma spray methods, with a total TBC layer thickness of 300-2000 microns (m).
- a ceramic insert 130 incorporates the profile of partially completed cooling passages, and optionally other types of engineered surface features (“ESFs”) 140 .
- ESFs are described in the aforementioned International Application No. PCT/US15/16318.
- the ceramic shell insert 130 is typically a partially sintered ceramic (similar to typical core material for investment casting processes) that is placed or positioned onto the wax pattern 150 .
- the ceramic shell insert 130 is placed into the wax injection tool or mold 142 and incorporated in the wax pattern 150 when it is injected and subsequently hardened.
- the wax pattern 150 (incorporating the ceramic insert 130 and a ceramic core 144 ) is dipped in and coated with ceramic slurry.
- the slurry is hardened, to form an investment casting outer shell mold 152 .
- the inner or interface surface 132 of the ceramic shell insert 130 (which is attached to the wax pattern 150 surface) incorporates the casting mold surface details for the eventual metal surfaces and structurally cast-in-place, cooling passage features.
- the cooling passage features are defined in the ceramic shell insert 130 by the projecting ceramic posts 134 that conform to the corresponding, partial cooling passage profiles.
- This casting method retains detail in the surface profile features, including the cooling passage profiles, which would otherwise be compromised in a wax pattern 150 due to fragility of the wax material composition.
- the ceramic shell insert 130 surface profile creation process for the superalloy component lends itself to modularity, where additional partially completed cooling passage forming ceramic posts 134 , and engineered surface feature anchoring surfaces 140 are incorporated for exposed airfoil areas such as leading edges and trailing edges of turbine blades 92 or vanes 104 , 106 .
- the ceramic shell inserts 130 are partially thermally processed prior to application to the wax injection tool 142 .
- the shell insert 130 remains as part of the outer ceramic shell structure 152 , which defines the outer cavity wall for the investment cast surfaces e.g., the turbine blade 92 , concave profile, high-pressure side 96 and the convex low-pressure side 98 surfaces.
- the ceramic shell insert system 130 exemplary embodiments of FIGS. 3-8 provide the ability to cast one or an array of engineered surface features and/or partial cooling hole passages 99 , 105 in a blade 92 or vane 104 , 106 within the engine 80 of FIG. 2 , through ESFs 140 and ceramic posts 134 , that are within the ceramic insert interface or inner surface 132 .
- a partial cooling passage is manufactured by creating a ceramic insert 130 with the ceramic rod protrusions or posts 134 .
- the ceramic posts 134 are integrally formed within, or formed separately and bonded to the inner surface 132 of the ceramic shell insert 130 .
- the ceramic posts or rods 134 penetrate, or in other words embed, within the wax pattern 150 , creating a partially complete cooling hole 99 or 105 .
- the ceramic shell insert 130 is not integral with an internal ceramic core 144 , and in exemplary embodiment herein, the projecting ceramic posts 134 do not contact the internal core, 144 , leaving an incomplete cooling passage.
- a partial cooling passage/hole beneficially reduces processing time needed to cut a complete cooling passage within a solid superalloy component after casting.
- casting superalloy, combustion turbine components, with partially completed or formed cooling passages 99 and 105 advantageously simplifies cooling passage completion, and reduces likelihood of previously applied thermal barrier coat delamination, during cooling passage completion.
- FIGS. 5-12 The main steps for investment casting of a combustion turbine component with partially completed cooling passages, in accordance with embodiments of the invention methods, are shown in FIGS. 5-12 . Each step is described generally as follows.
- one or more shell inserts 130 are provided to match desired surface profile of the engine component cooling passage profiles or engineered surface features.
- the completed ceramic shell insert 130 is oriented within a solid wax pattern 150 .
- the ceramic shell insert 130 is pushed directly into a previously completed, hardened wax pattern 150 that mimics the rest of the component profile.
- the ceramic insert 130 is positioned within an internal cavity 146 of a wax injection mold or die 142 , along with any other desired inserts, such as the ceramic core 144 , in a spaced relationship from other mold surfaces.
- the ceramic core corresponds to internal hollow portion surface profile of the engine component, such as a blade or vane cooling plenum.
- multiple, modular ceramic shell inserts 130 are used to form the entire desired surface of the component casting.
- Molten wax 148 is interposed or injected into the mold or die cavity 146 , which envelops therein the posts 134 of the ceramic shell insert, and any other types of engineered surface features. Gaps between the ceramic posts 134 and other mold surface features, such as a ceramic core 144 , are now filled with molten wax 148 , which subsequently hardens into a wax pattern 150 .
- the hardened wax pattern 150 which now captures the ceramic shell insert 130 , the posts 134 and the ceramic inner core 144 , is separated from the mold 142 , leaving the composite pattern of FIGS. 6 and 9 , which conforms to the outer profile of the desired engine component.
- the composite component pattern, including hardened wax pattern 150 , the shell insert 130 and ceramic core 144 are dipped or otherwise enveloped in ceramic slurry, dried, and fired in known types of investment casting manufacturing processes, to form a ceramic outer shell or casting vessel 152 , forming a casting mold for the superalloy component.
- the ceramic/wax composite vessel 152 and pattern 150 is dewaxed, such as in a known autoclave, leaving a composite ceramic vessel, with the hollow cavity 146 , shown in FIGS. 7 and 10 .
- the composite ceramic vessel 152 hollow cavity 146 incorporates the surface features of the superalloy component, including the partial cooling passages/holes.
- the post 134 does not contact the ceramic inner core 144 , leaving an open-space gap G that will be filled with superalloy material during the subsequent casting process.
- the total substrate wall thickness G of the cast component is established by the distance between the inner surface of the ceramic core 144 and the inner surface 132 , of the ceramic shell insert 130 , within the mold cavity 146 .
- the ceramic vessel 152 including the ceramic shell insert 130 and the ceramic core 144 , is filled with molten superalloy metal 154 , typically in a vacuum casting process.
- the post 134 including the post rim portion 136 and post tip 138 , will form the profile of a partially completed cooling passage having a partial cooling passage depth D, which as previously described is approximately 50% to 90% of the substrate wall thickness G.
- the remaining depth remnant of superalloy material 154 to remove from the component, in order to complete the cooling passage to the outer boundary of the component is G minus D.
- the now ceramic-free metal casting 154 now has a partially completed cooling passage 156 of partial depth D compared to the total substrate thickness G.
- the partially completed cooling passage 156 includes an entrance 158 and a terminus or hole bottom 160 , shown in FIG. 12 . In this embodiment, the terminus 160 is laterally offset from the passage entrance 160 .
- the partially completed cooling passage 156 is shown schematically as a cylindrical passage, oriented at an angle ⁇ relative to the component surface 161 .
- the partially completed cooling passages in the now ceramic-free superalloy metal casting are completed, before or after optional TBC application, by removing remnant metal from the casting by mechanical cutting, pressurized water or other fluid jet, or ablation processes.
- remnant metal removal to complete the cooling passage is initiated inside the partial cooling passage 156 , starting at the terminus 160 .
- remnant metal is removal is initiated from outside the substrate (here opposite the partially completed cooling passage entrance 158 ) until the newly created passage portion is in communication with the terminus of the partially completed cooling passage 156 .
- post 134 dimension definitions are summarized as follows.
- the posts 134 are oriented at an angle ⁇ of less than 90 degrees and typically 30 to 60 degrees relative to the ceramic shell insert surface 132 .
- the posts 134 have a diameter or thickness “t”, which is chosen to match the cross sectional diameter of a corresponding cooling passage.
- the post diameter t is typically 0.7 mm to 1.75 mm and the post 134 length “L” is typically the cooling passage partial depth D multiplied by the cosine ⁇ .
- the cooling passage partial depth D is 50% to 95% of the total substrate thickness “G”.
- the ceramic shell insert 130 is manufactured with an array of ceramic posts that are profiled to mimic integrally cast, partial cooling passages, or holes.
- known, cut cooling holes are 0.5-0.6 mm diameter, cylindrical in shape, and at 30-degree angle ⁇ to the surface.
- this type of passage or hole 156 can be integrally created in the casting, rather than fully cut in solid metal after the casting process, by using the ceramic post structure 134 in the ceramic shell insert 130 .
- geometry of the ceramic post 134 is not limited to simple shapes and angles.
- FIGS. 13-22 Various exemplary partial cooling passage profiles are shown in FIGS. 13-22 , which will be described in greater detail herein.
- the ceramic shell insert 130 has a plurality of ceramic posts 134 , corresponding to a pattern of partially completed or partial-depth cooling passages in the engine component. In some embodiments, a plurality of ceramic shell inserts 130 are used to form partial cooling passages within a component. In one or more embodiments, at least one ceramic post 134 projects from the ceramic insert surface 132 at an angle ⁇ , which is less than 90 degrees, and in other embodiments at least one ceramic post 134 projects from the ceramic insert surface 132 at an angle ⁇ of between 30 and 60 degrees.
- integrally cast, partial cooling passages formed by the method embodiments of the present invention, allow cooling fluid flow, heat transfer, and TBC delamination inhibiting design options that cannot be easily replicated by known post-casting cooling passage formation processes, with easier manufacture than passages formed by known refractory metal core (“RMC”) insert processes.
- Cooling passage/hole configurations are not limited to simple cylindrical holes, as shown in FIG. 12 . Changing the diameter or cross section or passage path of the cooling passage profile through the thickness of the component wall beneficially offers ability to increase or decrease the velocity of the cooling flow, depending on whether the hole diameter/cross section is decreased or increased as it approaches the outer surface of the component.
- FIGS. 13-22 show alternative cooling passage profiles that are formed in investment cast, superalloy components.
- the ceramic inner core 144 forms an internal surface of an engine component, such as a turbine blade or vane.
- FIGS. 13-18 show exemplary embodiments of change of cooling passage shape made by the casting methods of the present invention.
- changes in cooling passage shape have been previously made by EDM and laser ablation, as well as by other known material cutting methods.
- film-cooling passages on turbine blades and vanes have been previously made by EDM or millisecond laser drilling, and then a trapezoidal/pyramidal shape, flared outer surface profile is created on the surface by use of a finer nanosecond laser.
- FIG. 13-22 show alternative cooling passage profiles that are formed in investment cast, superalloy components.
- the ceramic inner core 144 forms an internal surface of an engine component, such as a turbine blade or vane.
- FIGS. 13-18 show exemplary embodiments of change of cooling passage shape made by the casting methods of the present invention.
- the ceramic post 164 has a diverging cylindrical profile toward the ceramic insert 162 , which will form a corresponding, divergent cooling passage profile in a turbine engine component outer surface.
- the ceramic post 168 has a converging cylindrical profile toward the ceramic insert 166 .
- the ceramic post 172 that is formed in ceramic insert 170 has a rectangular profile
- the ceramic post 176 of the ceramic insert 174 has a trapezoidal profile.
- a projecting, trapezoidal profile post 180 in the ceramic insert 178 has a smooth, angular profile transition to the post tip 181 , which corresponds to the terminus, or inner-most reaching portion or pinnacle of the partially completed cooling passage, whereas the post 184 on the ceramic insert 182 of FIG. 18 has a stepped transition from a trapezoidal entrance portion 186 to a necked, cylindrical portion 188 at the post tip or terminus.
- FIGS. 19 and 20 Additional embodiments for forming posts in ceramic inserts are shown in FIGS. 19 and 20 , where one cooling fluid feed from an exemplary blade or vane component-cooling cavity is split into many cooling ejections on the surface.
- the projecting post 200 of the ceramic insert 190 creates a cooling passage with a single feed plenum at the terminus 208 that splits into three cooling path outlets 202 , 204 , 206 , which will exhaust from the component surface.
- a cooling passage with more or less than three splits can be constructed in accordance with embodiments of the present invention.
- An alternative embodiment is shown in FIG.
- FIGS. 21-22 Such multi-dimensional, cooling passage formation method is applicable to two-dimensional or three-dimensional profiles within a component (into or out of the drawing figure sheet).
- Embodiments of the invention also fabricate non-linear cooling passage paths, which are shown in FIGS. 21-22 .
- the ceramic insert 220 of FIG. 21 includes a non-linear post 222 , with an entrance 224 and a terminus 226 .
- the cooling passage is completed after casting by removing material from the component, so that the terminus 226 extends outside the component wall boundary defined by the ceramic inner core 144 .
- FIG. 21 The ceramic insert 220 of FIG. 21 includes a non-linear post 222 , with an entrance 224 and a terminus 226 .
- the cooling passage is completed after casting by removing material from the component, so that the terminus 226 extends outside the component wall boundary defined by the ceramic inner core 144 .
- pigtail-like cooling passage is formed by post 230 , with an initial end portion 234 formed at the ceramic insert 228 , and extending to a terminus 236 towards the ceramic core insert 144 .
- the post 230 extends into a third dimension (into or out of the drawing figure sheet).
- Non-linear passages with serpentine paths, formed by the methods of the embodiments herein, provide for longer cooling paths within the same component volume, providing greater heat transfer surface area and longer dwell time for cooling flow, compared to a linear passage occupying the same volume in the component, at any given cooling flow velocity below the respective passage choke velocity.
- cooling passages formed in superalloy engine components before application of thermal barrier coating (“TBC”) layers are masked to inhibit obstruction by the later applied TBC material, which is costly and time consuming.
- TBC thermal barrier coating
- the substrate 240 has an over layer of bond coat (“BC”), which is typically MCrAlY material, before application of the TBC layer 244 .
- BC bond coat
- After application of the TBC layer 244 it is typically drilled with a laser 246 ablation device that uses percussions (shown schematically by the bubble 248 ) to pulse through the TBC 244 ( FIG.
- TBC thermal barrier coating
- the skewed path of the partially completed cooling passage 264 within the component substrate 262 interposes an overhanging shield layer of superalloy material 269 in the passage wall that is interposed between the laterally offset passage terminus 268 and the partial cooling passage entrance 266 that is proximate the TBC material in the zone 278 .
- the partially completed cooling passage 264 is formed by any previously known cutting/or ablation method within the component surface, but beneficially such partially completed cooling passages 264 are formed in some embodiments by use of the projecting ceramic post, ceramic inserts 130 of the type shown in FIGS. 3-22 herein, during investment casting of the component 260 . If a non-investment casting process is used to manufacture the engine component, the partially completed cooling passages 264 are likely formed by removing material from the component, prior to BC 274 and TBC layer(s) 276 application.
- the investment casting method embodiments and the TBC layer protection method embodiments are independent, but may be practiced jointly when fabricating investment cast, superalloy components for turbine engines, such as blades and vanes, having TBC layers.
- FIGS. 26 and 27 illustrate embodiments of methods for forming a cooling passage 270 in an investment cast, superalloy blade or vane component for a combustion turbine engine, where the component 260 has a component wall (e.g., the superalloy substrate 262 ) that is delimited by respective first and second wall surfaces (e.g., the respective upper surface at the junction with the bond coat 274 and the lower surface defining the cooling passage margin 272 ).
- first and second wall surfaces e.g., the respective upper surface at the junction with the bond coat 274 and the lower surface defining the cooling passage margin 272 .
- the completed cooling passage 270 will extend through the component wall (substrate 262 ) between its respective first and second surfaces, with the first surface forming an outer surface, for exposure to combustion gas, and the second surface forming an inner surface that is in communication with cooling channels formed in the blade or vane component 260 .
- the cooling passage formation methods to protect the TBC layer are applicable to other engine components, such as transitions 85 or combustor baskets within the engine 80 combustion section 84 .
- a partially completed cooling passage 264 is formed in a first surface of a wall of a superalloy engine component 260 for a combustion turbine engine.
- the partially formed or partially completed cooling passage 264 has an entrance 266 formed in the component substrate 262 first surface, which corresponds to a cooling passage inlet or outlet.
- the partially completed cooling passage 264 has a skewed passage path within the component wall substrate 262 , having a terminus 268 that is laterally offset from the passage entrance 266 , and distal the component first surface.
- the laterally offset passage entrance 266 and terminus 268 have an overhanging shield layer 269 of superalloy material in the wall that is interposed between the passage terminus 266 and the component first surface proximate the laterally offset passage entrance 266 .
- cooling passage 270 and the partially completed or partially formed passage 264 are shown in FIGS. 26 and 27 as having a straight path, symmetrical cross section, and constant skewed path angle in the component substrate 262 .
- the cooling passage defines a non-linear path, with or without an asymmetrical, axial or radial cross section. Examples of alternative embodiment cooling passage profiles are shown in FIGS. 13-22 .
- a thermal barrier coating 276 is applied over the component substrate 262 first surface and the partially completed or formed passage entrance 266 .
- the thermal barrier coating 276 comprises a known composition, thermally sprayed, or vapor deposited, or solution/suspension plasma sprayed thermal barrier coat that is applied directly to the component substrate 262 surface, or that is applied over an intermediate bond coat layer 274 that was previously applied over the component substrate surface.
- An ablation apparatus such as a pulsed laser 246 or an electric discharge machine, is used for ablating the thermal barrier coating 276 and the superalloy material in the substrate 262 .
- the laser 246 or other ablation device is aligned with the entrance 266 of the partially completed or formed passage 264 , and ablates thermal barrier coating material 276 from the partially completed or formed passage, reaching the passage terminus 268 .
- the cooling passage 270 is completed by ablating superalloy material out of the skewed path of the partially completed or formed passage 264 , along a cooling passage path 270 from the partial passage terminus 268 to the passage exit 272 on the second surface of the component substrate 262 , which forms the component 260 wall.
- the overhanging shield layer 269 of superalloy material in the partially formed/completed passage 269 inhibits damage to thermal barrier coating material 276 in the zone 278 proximate the passage entrance 266 .
- the otherwise avoided or mitigated TBC damage in the zone 278 is caused by ejection of ablated superalloy material out of the passage entrance 266 , which is shown schematically by the upwardly directed double arrow.
- a first level of ablation energy is applied with the ablation device 246 while ablating thermal barrier coating material from the partially completed or formed passage 264
- a second, higher level of ablation energy is applied with the ablation device 246 while ablating superalloy material in the substrate 262 , to complete the cooling passage 270 path.
- the first level of ablation energy applied at a lower pulse rate and/or lower energy intensity than the second level of ablation energy.
- the total ablation energy transferred to the component 260 during cooling passage 270 formation is insufficient to induce solidification cracking, or reheat cracking in the component during subsequent component heat treatment.
- EGFs engineered groove features 122 are formed in the TBC layer around part of or the entire periphery of turbine component cooling passages, such as cooling passages 99 of turbine blade 92 or cooling passages 105 of the vanes 104 and 106 , in order to limit delamination of the TBC over layer surrounding the cooling passage.
- the TBC layer at the extreme margin of the cooling passage entrance 99 or 105 , on the blade 92 or vanes 104 , 106 component surface can initiate separation from the metallic substrate that can spread laterally/horizontally within the TBC layer away from the hole.
- TBC delamination along one or more of the cooling hole 99 / 105 peripheral margins is arrested at the intersection of the circumscribing EGF segments 122 .
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Abstract
Description
- This application claims priority to International Application No. PCT/US15/16318, filed Feb. 18, 2015, and entitled “TURBINE COMPONENT THERMAL BARRIER COATING WITH CRACK ISOLATING ENGINEERED GROOVE FEATURES”, the entire contents of which are incorporated by reference herein.
- The invention relates to methods for forming cooling passages in superalloy substrates of combustion turbine components, such as engine blades, vanes, or transitions. More particularly, the invention relates to formation of partially completed cooling passages in such superalloy components during investment casting, by use of ceramic shell inserts, with projecting ceramic posts, then completing the passages in the castings.
- Known turbine engines, including gas/combustion turbine engines and steam turbine engines, incorporate shaft-mounted turbine blades circumferentially circumscribed by a turbine casing or housing. The remainder of this description focuses on applications within combustion or gas turbine technical application and environment, though exemplary embodiments described herein are applicable to steam turbine engines. In a gas/combustion turbine engine, hot combustion gasses flow in a combustion path that initiates within a combustor and are directed through a generally tubular transition into a turbine section. A forward or Row 1 vane directs the combustion gasses past successive alternating rows of turbine blades and vanes. Hot combustion gas striking the turbine blades cause blade rotation, thereby converting thermal energy within the hot gasses to mechanical work, which is available for powering rotating machinery, such as an electrical generator.
- Engine internal components within the hot combustion gas path are exposed to combustion temperatures approximately well over 1000 degrees Celsius (1832 degrees Fahrenheit). The engine internal components within the combustion path, such as for example combustion section transitions, vanes and blades are often constructed of high temperature resistant superalloys. Blades, vanes, and transitions often include cooling passages terminating in cooling holes on component outer surface, for passage of coolant fluid into the combustion path.
- Turbine engine internal components often incorporate a thermal barrier coat or coating (“TBC”) of metal-ceramic material that is applied directly to the external surface of the component substrate surface or over an intermediate metallic bond coat (“BC”) that was previously applied to the substrate surface. The TBC provides a thermal insulating layer over the component substrate, which reduces the substrate temperature. Combination of TBC application along with cooling passages in the component further lowers the substrate temperature.
- Fabrication of cooling passages in, and application of TBC layers to superalloy components, creates conflicting manufacturing constraints. Traditionally, cooling passages are formed by removing superalloy material from the intended passage path within the component, with exemplary removal tools including mechanical cutting/drilling bits, or various ablation devices, such as high-pressure water jet, percussion laser pulsation, and electric discharge machining (“EDM”). Cut cooling passage path, profile and size are limited by the physical capabilities of the cutting instrument. For example, drilled passages are linear and have cross sectional symmetry to match the drill bit. Ablated passages are limited by the size of the ablation instrument and ability to maneuver the instrument along a cutting path.
- Investment cast turbine engine components are fabricated by creating a hardened wax pattern, in a wax injection mold, which replicates the profile of the finished superalloy component. The wax pattern is enveloped in ceramic slurry, which is subsequently hardened by firing, into a ceramic shell casing. When wax is removed from the ceramic shell casing, the internal cavity is filled with molten superalloy material. Typically, more particularly, wax patterns for investment cast, superalloy components for combustion turbine engines, are injected into hard tool wax molds, and removed from the tools with precise and smooth surfaces. The wax patterns are then dipped in various ceramic slurry mixtures and processed to form the ceramic outer shell, which is subsequently sintered to form a vestibule in which molten metal is poured. Upon cooling and solidification, the outer ceramic shell is removed by mechanical and/or chemical methods and the metal part is then prepared for further processing. Further processing of the metal part includes ceramic core removal, finish machining, drilling of cooling holes, and application of a thermal barrier coating (“TBC”). Current state of the art processes often require the investment cast surface be lightly grit blasted to prepare the surface for bond coat application. At this point a bond coat, typically a metallic Cramoium, Aluminum, Yitria (“MCrAlY”) coating is applied to the substrate via a spray deposition technique, such as High Velocity Oxy Fuel (“HVOF”) or Low Pressure Plasma Spray (“LPPS”). After this a ceramic thermal barrier such as YSZ (Yttria Stabilized Zirconia) is applied to the surface of the MCrAlY via atmospheric or air plasma spray (“APS”) to complete the coating system. In some cases, a two layer ceramic coating is applied via APS for low thermal conductivity.
- The investment casting wax pattern does not have sufficient, reliable, structural integrity to form cooling passages directly therein. When cooling passages are formed in the mold that forms the wax pattern, there is more than insignificant chance that the cooling passage profile in the wax pattern will deform, or that the wax pattern passage will not fill completely with ceramic slurry; in either case the resultant passage in the metal casting does not confirm to design specification.
- In some investment casting, component manufacturing processes, refractory metal core (“RMC”) inserts that conform to the desired profiles and paths of cooling passages are placed in the molds prior to component metal casting. The RMC inserts have to be aligned precisely within the molds, and are removed after casting by chemical dissolution processes, adding to manufacturing complexity and expense. TBC layer application adds additional sequencing challenges to the manufacturing process.
- If cooling passages are formed in the blade, vane, transition, or other superalloy component prior to application of the TBC layer, the passages will become obstructed by the TBC material as the latter is applied to the component surface. Obstruction can be mitigated by temporarily masking the cooling passages on the component surface prior to the TBC application, which adds additional, costly, steps to the manufacturing processes. In the alternative, excess TBC material obstructions within cooling passages can be removed subsequently by the aforementioned cutting processes. Post TBC-application cooling passage obstruction removal increases risk of TBC layer damage and/or delamination along the margins of cooling passages on the component surface. In some manufacturing processes, cooling passages are formed after application of a TBC layer to the component substrate. In one known post TBC-coating cooling passage formation process, a pulsed laser ablates TBC material from the component at the intended cooling passage entry point, and then ablates the superalloy material to form the passage.
- As previously noted, there is risk of damage to the previously applied TBC layer, or delamination of the layer from the component substrate, as cooling passages are subsequently created within the component. Due to differences in thermal expansion, fracture toughness and elastic modulus, among other things, between typical metal-ceramic TBC materials and typical superalloy materials used to manufacture the aforementioned exemplary turbine components, there is potential risk of thermally- and/or mechanically-induced stress cracking of the TBC layer as well as TBC/turbine component adhesion loss at the interface of the dissimilar materials as the TBC layer and superalloy material are removed during cooling passage formation or cooling passage cleaning to remove TBC obstructions. The cracks and/or adhesion loss/delamination negatively affect the TBC layer's structural integrity and potentially lead to its spallation (i.e., separation of the TBC insulative material from the turbine component).
- In this invention a ceramic insert, which incorporates an engineered surface with one or more projecting posts, conforming to respective partially completed cooling passage(s), is incorporated into the casting mold. In some embodiments, the ceramic insert is a partially sintered, ceramic material (similar to typical core material used in investment casting processes) that is positioned onto and/or embedded within the wax pattern. Alternately, the ceramic shell insert is placed into the wax injection tool and embedded within the wax pattern when it is injected.
- Exemplary embodiments of the methods described herein form cooling passages in combustion turbine engine components, such as blades, vanes, or transitions, during investment casting, through use of ceramic shell inserts within the casting mold. Ceramic posts formed in the ceramic shell insert have profiles conforming to profiles of partially completed cooling passages. The shell insert surface, including its posts, is embedded in a hardened wax pattern, whose profile conforms to the component profile. The wax pattern and its embedded shell insert are enveloped in ceramic slurry that is hardened to form an outer ceramic shell. After removal of the hardened wax, the cavity within the outer ceramic shell is filled with melted superalloy material, and thereafter hardened to form a superalloy component casting. The ceramic outer shell, shell insert and any cores are removed from the casting, which now has cast-in-place, partial cooling passages formed therein. The cooling passages are completed by removing remaining superalloy material in the cooling passage path.
- Passage path profile is selectively varied easily by varying profile of the ceramic insert posts, which are affixed to the ceramic insert, and which do not need additional alignment effort within the casting molds, compared to standard, known RMC inserts. The partially completed cooling path passages are easily cleared of any subsequently applied TBC layer obstructions, prior to completion of the passages. The cooling passage is completed by removing the rest of the superalloy material from the casting to match a corresponding, completed cooling passage profile in the corresponding component.
- Exemplary embodiments of the invention feature a method for forming a cooling passage in an investment cast, superalloy component for a combustion turbine engine, by providing a wax injection mold defining a mold cavity, whose mold cavity surface conforms to a corresponding surface profile of a component for a combustion turbine engine. A ceramic shell insert is provided, having an insert surface profile and at least one ceramic post projecting from the inner surface, which in combination conform to a corresponding surface profile of a partially completed cooling passage in the engine component. The ceramic shell insert is inserted into the wax injection mold. The shell insert surface and each ceramic post forms part of the mold cavity surface, with each post projecting into the mold cavity. The mold cavity is filled with wax, enveloping each post therein. The wax is hardened, creating a wax pattern that embeds the ceramic shell and each post therein. The wax injection mold is removed. The hardened wax pattern, along with the insert surface and each ceramic post conforms to the component surface profile. The hardened wax pattern and embedded shell insert are enveloped in ceramic slurry. The ceramic slurry is fired, thereby hardening the slurry into an outer ceramic shell, which is joined to the ceramic shell insert, and eliminating the wax. The joined outer ceramic shell and ceramic shell insert define a shell internal cavity, which conforms to the engine component profile, including each partially completed cooling passage. The shell internal cavity is filled with molten superalloy material. The superalloy material is cooled, to form a component casting therein. The casting includes each partially completed cooling passage. The outer ceramic shell and ceramic insert, including each ceramic post, are removed from the casting, exposing each partially completed cooling passage. At least one cooling passage is completed by removing hardened, cast superalloy material from the casting to match a corresponding, completed cooling passage profile in the corresponding component.
- Other exemplary embodiments of the invention feature a method for forming a cooling passage in an investment cast, superalloy blade or vane component for a combustion turbine engine. The component has a component wall delimited by a wall outer surface, and a wall inner surface that defines a component cavity therein, with the cooling passage extending through the component wall between its respective outer and inner surfaces. A wax injection mold is provided, defining a mold cavity, whose mold cavity surface conforms to a corresponding profile of an outer surface of a hollow blade or vane component for a combustion turbine engine. A ceramic shell insert is provided, having an insert surface profile and at least one ceramic post projecting from the insert surface, which in combination conform to a corresponding surface profile of a partially completed cooling passage in the engine component wall. A ceramic core is provided, whose core outer surface conforms to a corresponding profile of a wall inner surface of the same component. The ceramic core is inserted into the mold cavity, and the ceramic shell insert is inserted into the mold, with the shell insert surface in opposed, spaced relationship with the ceramic core outer surface, forming a cavity void there between, which corresponds to a corresponding profile of the component wall respective outer and inner surfaces. The mold cavity void is filled with wax, enveloping each ceramic post therein. The wax is hardened, forming a wax pattern that embeds the ceramic core, and ceramic shell insert, including each ceramic post therein. The wax injection mold is removed, with the hardened wax pattern, along with the ceramic insert surface and each ceramic post conforming to the component wall respective outer and inner surfaces profiles. The hardened wax pattern and embedded shell insert are enveloped in ceramic slurry, which is fired, thereby hardening the slurry into an outer ceramic shell, which is joined to the ceramic shell insert and the ceramic core. The wax is eliminated, with the joined outer ceramic shell and ceramic insert defining a shell internal cavity, which conforms to the engine component wall outer surface profile, including each partially completed cooling passage. Similarly, the ceramic core outer surface conforms to the engine component wall inner surface profile, and the ceramic shell cavity void conforms to the engine component wall. The ceramic shell internal cavity void is filled with molten superalloy material, which is then cooled to form a component casting therein. The casting includes each partially completed cooling passage in the component outer wall surface. The ceramic core, ceramic outer shell, and ceramic insert, including each ceramic post, are removed from the casting, exposing each partially completed cooling passage in the outer surface of the component wall. At least one cooling passage is completed through the component wall, by removing hardened, cast superalloy material that blocks remaining path of the cooling passage.
- The respective features of the exemplary embodiments of the invention that are described herein may be applied jointly or severally in any combination or sub-combination.
- The exemplary embodiments of the invention are further described in the following detailed description in conjunction with the accompanying drawings, in which:
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FIG. 1 is a partial axial cross sectional view of a gas or combustion turbine engine incorporating one or more superalloy components, having cooling passages formed in accordance with exemplary method embodiments of the invention; -
FIG. 2 is a detailed cross sectional elevational view of the turbine engine ofFIG. 1 , showing Rows 1 turbine blade and Rows 1 and 2 vanes, having cooling passages formed in accordance with exemplary method embodiments of the invention; -
FIG. 3 is an elevational perspective view of an exemplary ceramic shell insert, having ceramic posts and engineered surface features (“ESFs”), which is used to form cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention; -
FIG. 4 is a perspective view of exemplary ceramic posts and ESFs of the ceramic shell insert ofFIG. 3 ; -
FIG. 5 is a cross-sectional plan view of a wax injection mold for a turbine blade airfoil, incorporating the ceramic shell insert ofFIG. 3 and a ceramic core, during wax injection into the mold cavity, which is used to form cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention; -
FIG. 6 is a cross-sectional plan view of a hardened wax pattern, with embedded ceramic shell insert and ceramic core, for a turbine blade airfoil, after removal from the wax injection mold ofFIG. 5 , which is used to form cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention; -
FIG. 7 is a cross-sectional plan view of the ceramic shell insert and ceramic core after envelopment in an outer ceramic shell and removal of the hardened wax, which is used to form cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention; -
FIG. 8 is a cross-sectional view of a ceramic shell insert and ceramic post, which is used to form cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention; -
FIG. 9 is a cross-sectional view of the ceramic shell insert and ceramic post ofFIG. 8 , in opposed orientation with the ceramic core, after wax injection, in accordance with exemplary method embodiments of the invention; -
FIG. 10 is a cross-sectional view of the ceramic shell insert and ceramic post ofFIG. 8 , in opposed orientation with the ceramic core, after envelopment in an outer ceramic shell and removal of the hardened wax, in accordance with exemplary method embodiments of the invention; -
FIG. 11 is a cross-sectional view of the ceramic shell insert and ceramic post ofFIG. 8 , in opposed orientation with the ceramic core, after filling the mold with molten superalloy material and subsequent hardening of the material, in accordance with exemplary method embodiments of the invention; -
FIG. 12 is a cross-sectional elevation view of the superalloy component casting, after removal of the ceramic shell insert, ceramic core and ceramic outer shell, and a formed, partial depth cooling passage formed therein, in accordance with exemplary method embodiments of the invention; -
FIG. 13 is a cross-sectional view of a ceramic shell insert and alternative embodiment converging frustro-conical profile ceramic post, in opposed orientation with a ceramic core, similar toFIG. 8 , which is used to form cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention; -
FIG. 14 is a cross-sectional view of a ceramic shell insert and alternative embodiment diverging profile ceramic post, in opposed orientation with a ceramic core, similar toFIG. 8 , which is used to form cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention; -
FIG. 15 is a cross-sectional view of a ceramic shell insert and alternative embodiment profile rectangular profile ceramic post, in opposed orientation with a ceramic core, similar toFIG. 8 , which is used to form cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention; -
FIG. 16 is a cross-sectional view of a ceramic shell insert and alternative embodiment trapezoidal profile ceramic post, in opposed orientation with a ceramic core, similar toFIG. 8 , which is used to form cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention; -
FIGS. 17 and 18 are cross-sectional views of ceramic shell inserts and alternative embodiment trapezoidal profile ceramic posts in opposed orientation with ceramic cores, similar toFIG. 8 , which are used to form film cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention; -
FIGS. 19 and 20 are cross-sectional views of ceramic shell inserts and alternative embodiment split-profile ceramic posts in opposed orientation with ceramic cores, similar toFIG. 8 , which are used to form split cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention; -
FIGS. 21 and 22 are cross-sectional views of ceramic shell inserts and alternative embodiment non-linear, asymmetrical profile ceramic posts in opposed orientation with ceramic cores, similar toFIG. 8 , which are used to non-linear and/or asymmetrical cooling passages in superalloy components, in accordance with exemplary method embodiments of the invention; -
FIGS. 23-25 are elevational cross section views of a prior art method for forming cooling passages in a thermal barrier coated (“TBC”) superalloy component for a combustion turbine engine, leading to undesired TBC delamination from the component substrate and its bond coat (“BC”) layer around the passage margins; and -
FIGS. 26 and 27 are elevational cross section views of an exemplary method for forming cooling passages in a thermal barrier coated (“TBC”) superalloy component for a combustion turbine engine, where a previously formed, partial cooling passage inhibits undesired TBC delamination from the component substrate and its bond coat (“BC”) layer around the passage margin, by shielding the TBC layer with an overhanging layer of superalloy material, in accordance with exemplary method embodiments of the invention. - To facilitate understanding, identical reference numerals have been used, where possible, to designate identical elements that are common to the figures. The figures are not drawn to scale.
- Exemplary method embodiments of the invention form partially completed cooling passages directly in turbine engine components during investment casting of the superalloy material. The partial cooling passages are formed by use of ceramic shell inserts, having projecting posts, and optionally additional engineered surface features. The ceramic posts have sufficient strength to maintain structural integrity and alignment as they are being embedded in wax during the wax pattern formation of the manufacturing process. After the wax hardens, the composite hardened wax pattern and embedded ceramic shell, along with any other casting mold segments, such as central cores, are in turn enveloped in ceramic slurry. The slurry is fired, hardening it into an outer ceramic shell. Any wax that was not previously eliminated during the outer ceramic shell firing is removed, leaving a mold cavity, for receipt of molten superalloy material. The ceramic posts of the ceramic shell insert project into the mold cavity. Molten superalloy material poured into the mold cavity hardens around the ceramic posts. After the casting cools and hardens, the ceramic mold material, including the outer shell, shell insert and the projecting ceramic posts are removed; leaving partially formed or completed cooling passages in the component substrate.
- In some embodiments, a TBC layer is applied over desired portions of the component substrate, including the partially completed cooling passages. Any TBC material obstructing the partial cooling passages is removed, allowing access to the terminus of the partial cooling passage. Thereafter the cooling passage path is completed by removing component superalloy material along the intended passage path.
- Referring to
FIGS. 1-2 , turbine engines, such as the gas orcombustion turbine engine 80 include amulti-stage compressor section 82, acombustion section 84, amulti-stage turbine section 86 and anexhaust system 88. Atmospheric pressure intake air is drawn into thecompressor section 82 generally in the direction of the flow arrows F along the axial length of theturbine engine 80. The intake air is progressively pressurized in thecompressor section 82 by rows rotating compressor blades and directed by mating compressor vanes to thecombustion section 84, where it is mixed with fuel and ignited. The ignited fuel/air mixture, now under greater pressure and velocity than the original intake air, is directed through atransition 85 to the sequential blade rows R1, R2, etc., in theturbine section 86. The engine's rotor andshaft 90 has a plurality of rows of airfoil cross sectional shapedturbine blades 92 terminating indistal blade tips 94 in thecompressor 82 andturbine 86 sections. - For convenience and brevity, further discussion of cooling passage formation and application of thermal barrier coat (“TBC”) layers on the combustion turbine engine components will focus on the
turbine section 86 embodiments and applications, though similar constructions are applicable for thecompressor 82 orcombustion 84 sections, as well as for steam turbine engine components. In the engine's 80turbine section 86, eachturbine blade 92 has a concave profile high-pressure side 96 and a convex low-pressure side 98. Coolingpassages 99 that are formed in theblade 92 facilitate passage of cooling fluid along the blade surface. The high velocity and pressure combustion gas, flowing in the combustion flow direction F imparts rotational motion on theblades 92, spinning therotor 90. As is well known, some of the mechanical power imparted on therotor shaft 90 is available for performing useful work. The combustion gasses are constrained radially distal therotor 90 byturbine casing 100 and proximal therotor 90 byair seals 102 comprising abradable surfaces. - Referring to the Row 1 section shown in
FIG. 2 , respectiveupstream vanes 104 anddownstream vanes 106 respectively direct upstream combustion gas generally parallel to the incident angle of the leading edge ofturbine blade 92 and redirect downstream combustion gas exiting the trailing edge of theblade 92 for a desired entry angle into downstream Row 2 turbine blades (not shown). Coolingpassages 105 that are formed in thevanes cooling passages FIG. 2 are merely schematic representations, are enlarged for visual clarity, and are not drawn to scale. Atypical turbine blade 92 orvane - As previously noted, turbine component surfaces that are exposed to combustion gasses are often constructed with a TBC layer for insulation of their underlying substrates. Typical TBC coated surfaces include the
turbine blades 92, thevanes ring segments 110,abradable surfaces 120 and related carrier surfaces of turbine vanes, and combustion section transitions 85. The TBC layer forblade 92,vanes ring segments 110, andtransition 85 exposed surfaces are often applied by thermal sprayed or vapor deposition or solution/suspension plasma spray methods, with a total TBC layer thickness of 300-2000 microns (m). - Fabrication of Partially Completed Cooling Passages with Ceramic Shell Inserts in Investment Cast, Engine Components
- Referring to
FIGS. 3-7 , in some embodiments of this invention, aceramic insert 130 incorporates the profile of partially completed cooling passages, and optionally other types of engineered surface features (“ESFs”) 140. ESFs are described in the aforementioned International Application No. PCT/US15/16318. Theceramic shell insert 130 is typically a partially sintered ceramic (similar to typical core material for investment casting processes) that is placed or positioned onto thewax pattern 150. Alternatively, as shown inFIG. 5 , theceramic shell insert 130 is placed into the wax injection tool ormold 142 and incorporated in thewax pattern 150 when it is injected and subsequently hardened. Thewax pattern 150, (incorporating theceramic insert 130 and a ceramic core 144) is dipped in and coated with ceramic slurry. The slurry is hardened, to form an investment castingouter shell mold 152. The inner orinterface surface 132 of the ceramic shell insert 130 (which is attached to thewax pattern 150 surface) incorporates the casting mold surface details for the eventual metal surfaces and structurally cast-in-place, cooling passage features. - The cooling passage features are defined in the
ceramic shell insert 130 by the projectingceramic posts 134 that conform to the corresponding, partial cooling passage profiles. This casting method retains detail in the surface profile features, including the cooling passage profiles, which would otherwise be compromised in awax pattern 150 due to fragility of the wax material composition. Theceramic shell insert 130 surface profile creation process for the superalloy component lends itself to modularity, where additional partially completed cooling passage formingceramic posts 134, and engineered surface feature anchoring surfaces 140 are incorporated for exposed airfoil areas such as leading edges and trailing edges ofturbine blades 92 orvanes outer casting shell 152 shrinkage, in some embodiments the ceramic shell inserts 130 are partially thermally processed prior to application to thewax injection tool 142. In the example of anengine vane blade 92 ofFIG. 2 , when the castingshell 152 is thermally treated, theshell insert 130 remains as part of the outerceramic shell structure 152, which defines the outer cavity wall for the investment cast surfaces e.g., theturbine blade 92, concave profile, high-pressure side 96 and the convex low-pressure side 98 surfaces. - The ceramic
shell insert system 130 exemplary embodiments ofFIGS. 3-8 provide the ability to cast one or an array of engineered surface features and/or partialcooling hole passages blade 92 orvane engine 80 ofFIG. 2 , throughESFs 140 andceramic posts 134, that are within the ceramic insert interface orinner surface 132. A partial cooling passage is manufactured by creating aceramic insert 130 with the ceramic rod protrusions or posts 134. Theceramic rods 134 ofFIGS. 4 and 8 have arim portion 136 that is formed to match the corresponding, intended inlet or outlet hole profile of a cooling passage within a combustion turbine component, and apost tip portion 138 whose surface profile and distal end define the corresponding surface profile of the terminus or end of the partially completed cooling passage. Theceramic posts 134 are integrally formed within, or formed separately and bonded to theinner surface 132 of theceramic shell insert 130. The ceramic posts orrods 134 penetrate, or in other words embed, within thewax pattern 150, creating a partiallycomplete cooling hole ceramic shell insert 130 is not integral with an internalceramic core 144, and in exemplary embodiment herein, the projectingceramic posts 134 do not contact the internal core, 144, leaving an incomplete cooling passage. However, a partial cooling passage/hole beneficially reduces processing time needed to cut a complete cooling passage within a solid superalloy component after casting. As will be described in detail herein, casting superalloy, combustion turbine components, with partially completed or formedcooling passages - The main steps for investment casting of a combustion turbine component with partially completed cooling passages, in accordance with embodiments of the invention methods, are shown in
FIGS. 5-12 . Each step is described generally as follows. - Referring to
FIGS. 5 and 8 , one or more shell inserts 130 are provided to match desired surface profile of the engine component cooling passage profiles or engineered surface features. The completedceramic shell insert 130 is oriented within asolid wax pattern 150. In some embodiments, theceramic shell insert 130 is pushed directly into a previously completed, hardenedwax pattern 150 that mimics the rest of the component profile. In other embodiments, due to potential fragility of theceramic posts 134, theceramic insert 130 is positioned within aninternal cavity 146 of a wax injection mold or die 142, along with any other desired inserts, such as theceramic core 144, in a spaced relationship from other mold surfaces. The ceramic core corresponds to internal hollow portion surface profile of the engine component, such as a blade or vane cooling plenum. In some embodiments, multiple, modular ceramic shell inserts 130 are used to form the entire desired surface of the component casting.Molten wax 148 is interposed or injected into the mold or diecavity 146, which envelops therein theposts 134 of the ceramic shell insert, and any other types of engineered surface features. Gaps between theceramic posts 134 and other mold surface features, such as aceramic core 144, are now filled withmolten wax 148, which subsequently hardens into awax pattern 150. - The hardened
wax pattern 150, which now captures theceramic shell insert 130, theposts 134 and the ceramicinner core 144, is separated from themold 142, leaving the composite pattern ofFIGS. 6 and 9 , which conforms to the outer profile of the desired engine component. The composite component pattern, includinghardened wax pattern 150, theshell insert 130 andceramic core 144 are dipped or otherwise enveloped in ceramic slurry, dried, and fired in known types of investment casting manufacturing processes, to form a ceramic outer shell or castingvessel 152, forming a casting mold for the superalloy component. The ceramic/waxcomposite vessel 152 andpattern 150 is dewaxed, such as in a known autoclave, leaving a composite ceramic vessel, with thehollow cavity 146, shown inFIGS. 7 and 10 . - The composite
ceramic vessel 152hollow cavity 146 incorporates the surface features of the superalloy component, including the partial cooling passages/holes. In the embodiment of both ofFIGS. 7 and 10 , thepost 134 does not contact the ceramicinner core 144, leaving an open-space gap G that will be filled with superalloy material during the subsequent casting process. The total substrate wall thickness G of the cast component is established by the distance between the inner surface of theceramic core 144 and theinner surface 132, of theceramic shell insert 130, within themold cavity 146. InFIG. 11 , theceramic vessel 152, including theceramic shell insert 130 and theceramic core 144, is filled withmolten superalloy metal 154, typically in a vacuum casting process. Ultimately thepost 134, including thepost rim portion 136 andpost tip 138, will form the profile of a partially completed cooling passage having a partial cooling passage depth D, which as previously described is approximately 50% to 90% of the substrate wall thickness G. The remaining depth remnant ofsuperalloy material 154 to remove from the component, in order to complete the cooling passage to the outer boundary of the component is G minus D. - In
FIG. 11 , casting of thesuperalloy material 154 in the outer ceramic shell-castingmold 152 is completed, trapping thepost 134, including therim 136 andpost tip portion 138 in the now hardened metal. Once thesuperalloy metal 154 has solidified and cooled, ceramic mold material forming theformer shell insert 130 andposts 134,ceramic core 144, and the outerceramic shell 152 ofFIG. 7 , is removed from the hardened metal casting 154 mechanically and chemically. Typical, known chemical removal processes for ceramic mold material use heated sodium hydroxide (NaOH), and/or potassium hydroxide (KOH) baths. - The now ceramic-free metal casting 154 now has a partially completed cooling
passage 156 of partial depth D compared to the total substrate thickness G. The partially completedcooling passage 156 includes anentrance 158 and a terminus orhole bottom 160, shown inFIG. 12 . In this embodiment, theterminus 160 is laterally offset from thepassage entrance 160. The partially completedcooling passage 156 is shown schematically as a cylindrical passage, oriented at an angle θ relative to thecomponent surface 161. The partially completed cooling passages in the now ceramic-free superalloy metal casting are completed, before or after optional TBC application, by removing remnant metal from the casting by mechanical cutting, pressurized water or other fluid jet, or ablation processes. In some embodiments, remnant metal removal to complete the cooling passage is initiated inside thepartial cooling passage 156, starting at theterminus 160. Alternatively, remnant metal is removal is initiated from outside the substrate (here opposite the partially completed cooling passage entrance 158) until the newly created passage portion is in communication with the terminus of the partially completed coolingpassage 156. - Referring to
FIGS. 8, 10 and 12 , post 134 dimension definitions are summarized as follows. In some embodiments, theposts 134 are oriented at an angle θ of less than 90 degrees and typically 30 to 60 degrees relative to the ceramicshell insert surface 132. Theposts 134 have a diameter or thickness “t”, which is chosen to match the cross sectional diameter of a corresponding cooling passage. In embodiments herein, the post diameter t is typically 0.7 mm to 1.75 mm and thepost 134 length “L” is typically the cooling passage partial depth D multiplied by the cosine θ. In some embodiments, the cooling passage partial depth D is 50% to 95% of the total substrate thickness “G”. - As previously described, the
ceramic shell insert 130 is manufactured with an array of ceramic posts that are profiled to mimic integrally cast, partial cooling passages, or holes. Typically, known, cut cooling holes, not formed by the methods of this invention, are 0.5-0.6 mm diameter, cylindrical in shape, and at 30-degree angle θ to the surface. As shown inFIG. 12 , this type of passage orhole 156 can be integrally created in the casting, rather than fully cut in solid metal after the casting process, by using theceramic post structure 134 in theceramic shell insert 130. However, geometry of theceramic post 134 is not limited to simple shapes and angles. Various exemplary partial cooling passage profiles are shown inFIGS. 13-22 , which will be described in greater detail herein. In one or more embodiments, theceramic shell insert 130 has a plurality ofceramic posts 134, corresponding to a pattern of partially completed or partial-depth cooling passages in the engine component. In some embodiments, a plurality of ceramic shell inserts 130 are used to form partial cooling passages within a component. In one or more embodiments, at least oneceramic post 134 projects from theceramic insert surface 132 at an angle θ, which is less than 90 degrees, and in other embodiments at least oneceramic post 134 projects from theceramic insert surface 132 at an angle θ of between 30 and 60 degrees. - As previously noted, integrally cast, partial cooling passages, formed by the method embodiments of the present invention, allow cooling fluid flow, heat transfer, and TBC delamination inhibiting design options that cannot be easily replicated by known post-casting cooling passage formation processes, with easier manufacture than passages formed by known refractory metal core (“RMC”) insert processes. Cooling passage/hole configurations are not limited to simple cylindrical holes, as shown in
FIG. 12 . Changing the diameter or cross section or passage path of the cooling passage profile through the thickness of the component wall beneficially offers ability to increase or decrease the velocity of the cooling flow, depending on whether the hole diameter/cross section is decreased or increased as it approaches the outer surface of the component. -
FIGS. 13-22 show alternative cooling passage profiles that are formed in investment cast, superalloy components. In all ofFIGS. 13-22 , the ceramicinner core 144 forms an internal surface of an engine component, such as a turbine blade or vane.FIGS. 13-18 show exemplary embodiments of change of cooling passage shape made by the casting methods of the present invention. By way of background, changes in cooling passage shape have been previously made by EDM and laser ablation, as well as by other known material cutting methods. For example, film-cooling passages on turbine blades and vanes have been previously made by EDM or millisecond laser drilling, and then a trapezoidal/pyramidal shape, flared outer surface profile is created on the surface by use of a finer nanosecond laser. InFIG. 13 , theceramic post 164 has a diverging cylindrical profile toward theceramic insert 162, which will form a corresponding, divergent cooling passage profile in a turbine engine component outer surface. Conversely, inFIG. 14 , theceramic post 168 has a converging cylindrical profile toward theceramic insert 166. InFIG. 15 , theceramic post 172 that is formed inceramic insert 170 has a rectangular profile, whereas inFIG. 16 , theceramic post 176 of theceramic insert 174 has a trapezoidal profile. InFIG. 17 , a projecting,trapezoidal profile post 180 in theceramic insert 178 has a smooth, angular profile transition to the post tip 181, which corresponds to the terminus, or inner-most reaching portion or pinnacle of the partially completed cooling passage, whereas thepost 184 on theceramic insert 182 ofFIG. 18 has a stepped transition from atrapezoidal entrance portion 186 to a necked,cylindrical portion 188 at the post tip or terminus. - Additional embodiments for forming posts in ceramic inserts are shown in
FIGS. 19 and 20 , where one cooling fluid feed from an exemplary blade or vane component-cooling cavity is split into many cooling ejections on the surface. InFIG. 19 , the projectingpost 200 of theceramic insert 190 creates a cooling passage with a single feed plenum at theterminus 208 that splits into three coolingpath outlets FIG. 20 , where the projectingpost 212, formed on theceramic insert 210, forms one passage at theterminus 218 that splits linearly into two passages/holes outlets FIGS. 21-22 . Theceramic insert 220 ofFIG. 21 includes anon-linear post 222, with anentrance 224 and aterminus 226. The cooling passage is completed after casting by removing material from the component, so that theterminus 226 extends outside the component wall boundary defined by the ceramicinner core 144. InFIG. 22 , pigtail-like cooling passage is formed bypost 230, with an initial end portion 234 formed at theceramic insert 228, and extending to aterminus 236 towards theceramic core insert 144. Thepost 230 extends into a third dimension (into or out of the drawing figure sheet). Non-linear passages with serpentine paths, formed by the methods of the embodiments herein, provide for longer cooling paths within the same component volume, providing greater heat transfer surface area and longer dwell time for cooling flow, compared to a linear passage occupying the same volume in the component, at any given cooling flow velocity below the respective passage choke velocity. - As previously noted, cooling passages formed in superalloy engine components before application of thermal barrier coating (“TBC”) layers are masked to inhibit obstruction by the later applied TBC material, which is costly and time consuming. Often in the past cooling passages have been formed in
superalloy engine components 239, after TBC layer application by laser ablation, such as shown in progression ofFIGS. 23-25 . Thesubstrate 240 has an over layer of bond coat (“BC”), which is typically MCrAlY material, before application of theTBC layer 244. After application of theTBC layer 244, it is typically drilled with alaser 246 ablation device that uses percussions (shown schematically by the bubble 248) to pulse through the TBC 244 (FIG. 23 ), then the BC bond coat 242 (FIG. 24 ), then thesubstrate 240 superalloy base material to create a cooling passage/hole (FIG. 25 ). Delamination and/or cracking damage is commonly observed in theTBC layer 244, at its interface with thebond coat 242, as a result of themultiple percussions 248 exerting a force on the overhanging TBC in the zone labelled 250, caused by drilling through thebase material 240, with thelaser ablation device 246. - Potential damage to thermal barrier coating (“TBC”) layer(s) 276 during
subsequent cooling passage 270 formation is mitigated by creation of a partially completed or formedcooling passage 264 in the superalloy,turbine engine component 260, and prior to application of theTBC layer 276 on the same surface, as shown inFIGS. 26 and 27 . There isless substrate 262 material (and resultingless percussions 248 generated by the laser ablation device 246) needed to drill through the remnant superalloy material below theterminus 268 of the partially completed coolingpassage 264, thereby reducing opportunity to cause repetitive percussion damage to the TBC layer at thezone 278. The skewed path of the partially completed coolingpassage 264 within thecomponent substrate 262 interposes an overhanging shield layer ofsuperalloy material 269 in the passage wall that is interposed between the laterally offsetpassage terminus 268 and the partialcooling passage entrance 266 that is proximate the TBC material in thezone 278. - In practicing the TBC damage mitigation method of embodiments of the invention, the partially completed cooling
passage 264 is formed by any previously known cutting/or ablation method within the component surface, but beneficially such partially completed coolingpassages 264 are formed in some embodiments by use of the projecting ceramic post,ceramic inserts 130 of the type shown inFIGS. 3-22 herein, during investment casting of thecomponent 260. If a non-investment casting process is used to manufacture the engine component, the partially completed coolingpassages 264 are likely formed by removing material from the component, prior toBC 274 and TBC layer(s) 276 application. The investment casting method embodiments and the TBC layer protection method embodiments are independent, but may be practiced jointly when fabricating investment cast, superalloy components for turbine engines, such as blades and vanes, having TBC layers. -
FIGS. 26 and 27 illustrate embodiments of methods for forming acooling passage 270 in an investment cast, superalloy blade or vane component for a combustion turbine engine, where thecomponent 260 has a component wall (e.g., the superalloy substrate 262) that is delimited by respective first and second wall surfaces (e.g., the respective upper surface at the junction with thebond coat 274 and the lower surface defining the cooling passage margin 272). In these exemplaryFIGS. 26 and 27 the completedcooling passage 270 will extend through the component wall (substrate 262) between its respective first and second surfaces, with the first surface forming an outer surface, for exposure to combustion gas, and the second surface forming an inner surface that is in communication with cooling channels formed in the blade orvane component 260. It is noted that the cooling passage formation methods to protect the TBC layer are applicable to other engine components, such astransitions 85 or combustor baskets within theengine 80combustion section 84. - A partially completed cooling
passage 264 is formed in a first surface of a wall of asuperalloy engine component 260 for a combustion turbine engine. The partially formed or partially completed coolingpassage 264 has anentrance 266 formed in thecomponent substrate 262 first surface, which corresponds to a cooling passage inlet or outlet. The partially completedcooling passage 264 has a skewed passage path within thecomponent wall substrate 262, having aterminus 268 that is laterally offset from thepassage entrance 266, and distal the component first surface. The laterally offsetpassage entrance 266 andterminus 268 have an overhangingshield layer 269 of superalloy material in the wall that is interposed between thepassage terminus 266 and the component first surface proximate the laterally offsetpassage entrance 266. While thecooling passage 270 and the partially completed or partially formedpassage 264 are shown inFIGS. 26 and 27 as having a straight path, symmetrical cross section, and constant skewed path angle in thecomponent substrate 262. In other embodiments, the cooling passage defines a non-linear path, with or without an asymmetrical, axial or radial cross section. Examples of alternative embodiment cooling passage profiles are shown inFIGS. 13-22 . - A
thermal barrier coating 276 is applied over thecomponent substrate 262 first surface and the partially completed or formedpassage entrance 266. Thethermal barrier coating 276 comprises a known composition, thermally sprayed, or vapor deposited, or solution/suspension plasma sprayed thermal barrier coat that is applied directly to thecomponent substrate 262 surface, or that is applied over an intermediatebond coat layer 274 that was previously applied over the component substrate surface. - An ablation apparatus, such as a
pulsed laser 246 or an electric discharge machine, is used for ablating thethermal barrier coating 276 and the superalloy material in thesubstrate 262. Thelaser 246 or other ablation device is aligned with theentrance 266 of the partially completed or formedpassage 264, and ablates thermalbarrier coating material 276 from the partially completed or formed passage, reaching thepassage terminus 268. - In
FIG. 27 , thecooling passage 270 is completed by ablating superalloy material out of the skewed path of the partially completed or formedpassage 264, along acooling passage path 270 from thepartial passage terminus 268 to thepassage exit 272 on the second surface of thecomponent substrate 262, which forms thecomponent 260 wall. As previously noted, the overhangingshield layer 269 of superalloy material in the partially formed/completedpassage 269 inhibits damage to thermalbarrier coating material 276 in thezone 278 proximate thepassage entrance 266. The otherwise avoided or mitigated TBC damage in thezone 278 is caused by ejection of ablated superalloy material out of thepassage entrance 266, which is shown schematically by the upwardly directed double arrow. In some embodiments, a first level of ablation energy is applied with theablation device 246 while ablating thermal barrier coating material from the partially completed or formedpassage 264, and a second, higher level of ablation energy is applied with theablation device 246 while ablating superalloy material in thesubstrate 262, to complete thecooling passage 270 path. In some embodiments, the first level of ablation energy applied at a lower pulse rate and/or lower energy intensity than the second level of ablation energy. In some embodiments, the total ablation energy transferred to thecomponent 260 during coolingpassage 270 formation is insufficient to induce solidification cracking, or reheat cracking in the component during subsequent component heat treatment. - As described in the aforementioned, International Application No. PCT/US15/16318, filed Feb. 18, 2015, and entitled “TURBINE COMPONENT THERMAL BARRIER COATING WITH CRACK ISOLATING ENGINEERED GROOVE FEATURES”, in some embodiments, referring to
FIG. 2 , additional and optional engineered groove features (“EGFs”) 122 are formed in the TBC layer around part of or the entire periphery of turbine component cooling passages, such ascooling passages 99 ofturbine blade 92 orcooling passages 105 of thevanes cooling passage entrance blade 92 orvanes EGF 122 at a laterally spaced distance from thecooling hole cooling hole 99/105 peripheral margins is arrested at the intersection of the circumscribingEGF segments 122. - Although various embodiments that incorporate the invention have been shown and described in detail herein, others can readily devise many other varied embodiments that still incorporate the claimed invention. The invention is not limited in its application to the exemplary embodiment details of construction and the arrangement of components set forth in the description or illustrated in the drawings. The invention is capable of other embodiments and of being practiced or of being carried out in various ways. Also, it is to be understood that the phraseology and terminology used herein is for the purpose of description and should not be regarded as limiting. The use of “including,” “comprising,” or “having” and variations thereof herein is meant to encompass the items listed thereafter and equivalents thereof as well as additional items. Unless specified or limited otherwise, the terms “mounted”, “connected”, “supported”, and “coupled” and variations thereof are used broadly and encompass direct and indirect mountings, connections, supports, and couplings. Further, “connected” and “coupled” are not restricted to physical, mechanical, or electrical connections or couplings.
Claims (20)
Applications Claiming Priority (4)
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US14/188,941 US8939706B1 (en) | 2014-02-25 | 2014-02-25 | Turbine abradable layer with progressive wear zone having a frangible or pixelated nib surface |
US14/188,958 US9151175B2 (en) | 2014-02-25 | 2014-02-25 | Turbine abradable layer with progressive wear zone multi level ridge arrays |
PCT/US2015/016318 WO2015130526A2 (en) | 2014-02-25 | 2015-02-18 | Turbine component thermal barrier coating with crack isolating engineered groove features |
PCT/US2016/018215 WO2016133987A2 (en) | 2015-02-18 | 2016-02-17 | Forming cooling passages in combustion turbine superalloy castings |
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US15/548,267 Abandoned US20180015536A1 (en) | 2014-02-25 | 2016-02-17 | Forming cooling passages in combustion turbine superalloy castings |
US16/076,922 Abandoned US20190048730A1 (en) | 2014-02-25 | 2016-05-10 | Ceramic matrix composite turbine component with graded fiber-reinforced ceramic substrate |
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CN112912592A (en) * | 2018-07-13 | 2021-06-04 | 西门子能源全球两合公司 | Airfoils for turbine engines incorporating pins |
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KR102494689B1 (en) | 2018-07-13 | 2023-01-31 | 지멘스 에너지 글로벌 게엠베하 운트 코. 카게 | Airfoils for turbine engines incorporating fins |
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