WO2014113184A1 - Procédé permettant de former des orifices de refroidissement coulés dans un composant d'aéronef - Google Patents

Procédé permettant de former des orifices de refroidissement coulés dans un composant d'aéronef Download PDF

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Publication number
WO2014113184A1
WO2014113184A1 PCT/US2013/076906 US2013076906W WO2014113184A1 WO 2014113184 A1 WO2014113184 A1 WO 2014113184A1 US 2013076906 W US2013076906 W US 2013076906W WO 2014113184 A1 WO2014113184 A1 WO 2014113184A1
Authority
WO
WIPO (PCT)
Prior art keywords
wax
ceramic
core
projections
forming
Prior art date
Application number
PCT/US2013/076906
Other languages
English (en)
Inventor
Douglas Gerard Konitzer
James Herbert DELNES
Original Assignee
General Electric Company
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by General Electric Company filed Critical General Electric Company
Publication of WO2014113184A1 publication Critical patent/WO2014113184A1/fr

Links

Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B22—CASTING; POWDER METALLURGY
    • B22C—FOUNDRY MOULDING
    • B22C7/00—Patterns; Manufacture thereof so far as not provided for in other classes
    • B22C7/02—Lost patterns
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B22—CASTING; POWDER METALLURGY
    • B22C—FOUNDRY MOULDING
    • B22C9/00—Moulds or cores; Moulding processes
    • B22C9/02—Sand moulds or like moulds for shaped castings
    • B22C9/04—Use of lost patterns
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B22—CASTING; POWDER METALLURGY
    • B22C—FOUNDRY MOULDING
    • B22C9/00—Moulds or cores; Moulding processes
    • B22C9/10—Cores; Manufacture or installation of cores
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/20—Manufacture essentially without removing material
    • F05D2230/21—Manufacture essentially without removing material by casting
    • 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 disclosed embodiments generally pertain to one or more methods forming an aircraft engine component. More particularly, but not by way of limitation, present embodiments relate to a method of forming cast-in cooling holes in an aircraft engine component.
  • a typical gas turbine engine generally possesses a forward end and an aft end with its several core or propulsion components positioned axially therebetween.
  • An air inlet or intake is at a forward end of the engine. Moving toward the aft end, in order, the intake is followed by a compressor, a combustion chamber, a turbine, and a nozzle at the aft end of the engine, it will be readily apparent from those skilled in the art that additional components may also be included in the engine, such as, for example, low-pressure and high-pressure compressors, and high- pressure and low-pressure turbines. This, however, is not an exhaustive list.
  • An engine also typically has an internal shaft axially disposed along a center longitudinal axis of the engine. The internal shaft is connected to both the turbine and the air compressor, such that the turbine provides a rotational input to the air compressor to drive the compressor blades.
  • a high pressure turbine first receives the hot combustion gases from the combustor and includes a stator nozzle assembly directing the combustion gases downstream through a row of high pressure turbine rotor blades extending radially outwardly from a supporting rotor disk.
  • a second stage stator nozzle assembly is positioned downstream of the first stage blades followed in turn by a row of second stage rotor blades extending radially outwardly from a second supporting rotor disk.
  • cooling holes are utilized in order to maintain proper operating temperature of the part so that the part or component does not deteriorate or become damaged in the high temperature, pressure and stress environment of aircraft aviation or power generation. These cooling holes receive bypass or cooling air within the aircraft engine to pass through the parts or components and provide the cooling necessary for operation in the extreme conditions.
  • Current cooling holes are formed by machining the holes into the component after the component has been cast. This adds cost and time to the process of forming the components. Also, the process of making the cooling hole is somewhat blind and can be problematic because a core may shift during the casting process which causes variation in the actual cooling path location to which the cooling hole is machined.
  • a method of forming cast-in cooling holes wherein the cast-in holes are formed by projections disposed on a core. The projections are formed in a cast part and a portion of the cast metal is removed to reveal the cooling hole in the part.
  • a method of forming a cast-in cooling hole in an aircraft engine component comprises positioning a core in a die, the core having a projection corresponding to a recess in said die to aid in forming the cooling hole, injecting the die with wax about the core to form a wax pattern, the wax pattern having a wax projection corresponding to a recess in the die, coating the pattern with ceramic to form a mold having an outer shell, forming a ceramic projection along an inner surface of the mold, the ceramic projection corresponding to the wax projection, de- waxing the pattern, injecting the mold with molten metal, removing the ceramic shell and the core, and, removing a metal protuberance to reveal the cooling hole.
  • the method wherein the pattern is disposed on a wax tree assembly.
  • the method wherein the wax tree assembly is placed in a slurry.
  • the method further comprises applying dry ceramic pattern to the wax tree assembly after the slurry,
  • the method wherein the de-waxing comprises melting of the wax.
  • the method wherein the removing of the ceramic shell is performed with a chemical.
  • the method wherein the removing is a cutting step.
  • the method wherein the removing is a machining step.
  • the method wherein the removing is a grinding step.
  • the method further comprises forming an air foil.
  • the method further comprises forming a turbine blade.
  • the method farther comprises forming a stator vane.
  • the method further comprises forming a shroud.
  • a method of casting cooling holes in an aircraft engine component comprises positioning a core in a die, the core having projections corresponding to the cooling holes, positioning wax in the die and about the core, the core projections forming wax projections, coating the wax with ceramic-based material, the wax projections allowing formation of ceramic projections at least along an inner surface of ceramic shell, de-waxing the ceramic shell, flowing molten metal into the ceramic shell and about the ceramic projections, removing the ceramic shell, removing metal which surrounded the ceramic projections to reveal the cooling holes.
  • FIG. 1 is a side section view of a gas turbine engine.
  • FIG. 2 is a schematic view a core for the casting process.
  • FIG. 3 is a schematic view of the injection of wax into the die
  • FIG. 4 is a schematic view of the forming of a wax pattern.
  • FIG. 5 is a schematic view of the building of a tree of wax patterns.
  • FIG. 6 is a schematic view of the bathing of the wax pattern tree in an agent.
  • FIG. 7 is a schematic view of the stuccoing of the tree of wax patterns with ceramic.
  • FIG. 8 is a schematic view of the de- waxing of the tree of wax patterns.
  • FIG. 9 is a schematic view of the bum out of residue wax.
  • FIG. 30 is a schematic view of the pouring of molten metal into the remaining ceramic shell
  • FIG. ⁇ is a schematic view of the removal of the stucco shell.
  • FIG. 12 is a schematic view of the removal of parts from the tree.
  • FIG. 13 is a schematic view of the leaching out of ceramic material from the component tree.
  • FIG. 14 is a schematic view of the surface grinding of parts to reveal the cooling holes.
  • FIG. 15 is a side section view of a projection of ceramic material with the metallic overlay of the cast component.
  • FIG. 16 is a side section view of the component after the surface finishing reveals the cooling aperture.
  • FIG- 7 is a flow chart of certain steps of the method. DETAILED DESCRIPTION
  • cast-in cooling holes are formed by- use of a core having projections and a die having dimples or recesses to receive the projection to form a wax pattern.
  • the projection into the ceramic shell locates the cast-in cooling hole in the final cast metallic component.
  • the component is finished in any of various manners to reveal the cast-in cooling hole once the ceramic is removed from the component.
  • axial or axially refer to a dimension along a longitudinal axis of an engine.
  • forward used in conjunction with “axial” or “axially” refers to moving in a direction toward the engine inlet, or a component being relatively closer to the engine inlet as compared to another component.
  • aft used in conjunction with “axial” or “axialSy” refers to moving in a direction toward the engine nozzle, or a component being relatively closer to the engine nozzle as compared to another component.
  • the terms “radial” or “radially” refer to a dimension extending between a center longitudinal axis of the engine and an outer engine circumference.
  • proximal or “proximally,” either by themselves or in conjunction with the terms “radial” or “radially,” refers to moving in a direction toward the center longitudinal axis, or a component being relatively closer to the center longitudinal axis as compared to another component.
  • distal or disally, either by themselves or in conjunction with the terms “radial” or “radially,” refers to moving in a direction toward the outer engine circumference, or a component being relatively closer to the outer engine circumference as compared to another component.
  • lateral refers to a dimension that is perpendicular to both the axial and radial dimensions.
  • FIG. ⁇ a schematic side section view of a gas turbine engine 10 is shown having an engine inlet end 12 wherein air enters the propuisor or core 13 which is defined generally by a compressor 14, a combustor 16 and a multi-stage high pressure turbine 20. Collectively, the propuisor 13 provides thrust or power during operation.
  • the gas turbine 10 may be used for aviation, power generation, industrial, marine or the like.
  • air enters through the air inlet end 12 of the engine 10 and moves through at least one stage of compression where the air pressure is increased and directed to the eombustor 16. The compressed air is mixed with fuel and burned providing the hot combustion gas which exits the combustor 16 toward the high pressure turbine 20.
  • the shaft 24 passes toward the front of the engine to continue rotation of the one or more compressor stages 14, a turbofan 18 or inlet fan blades, depending on the turbine design.
  • the turbofan 18 is connected by the shaft 28 to a low pressure turbine 23 and creates thrust for the turbine engine 10.
  • a low pressure turbine 23 may also be utilized to extract further energy and power additional compressor stages.
  • the high pressure air may be used to aid in cooling components of the engine as well.
  • the gas turbine 10 is axis-symmetrical about engine axis 26 or shaft 24 so that various engine components rotate thereabout.
  • the axis-symmetrical shaft 24 extends through the turbine engine forward end into an aft end and is joumaled by bearings along the length of the shaft structure.
  • the shaft rotates about a centerline 26 of the engine 10.
  • the shaft 24 may be hollow to allow rotation of a low pressure turbine shaft 28 therein and independent of the shaft 24 rotation.
  • Shafts 28 also may rotate about the centerline axis 26 of the engine. During operation the shaft 28 rotates along with other structures connected to the shaft such as the rotor assemblies of the turbine in order to create power or thrust for various types of turbines used in power and industrial or aviation areas of use,
  • the core 10 may be formed in a variety manners.
  • the core 10 may formed by injection, casting or a variety of processes, including but not limited to additive manufacturing, of ceramic materials.
  • the core may be formed of aluminum, silicon, or zirconium based material. However, other materials may be utilized.
  • the core 10 includes projections 32 which will provide for flow through of cooling air in the cast component 60 (FIG. 1 1).
  • the core 10 is cured to form a hardened structure about which a wax pattern may be formed.
  • the core 10 is positioned within a die 34.
  • the projections 32 are received by dimples 36 within the die 34 so as to aid in positioning and locating the core 10,
  • This alignment or registration between the dimples 36 and the projections 32 provides two functions.
  • the dimples 36 help ensure the cooling holes are properly positioned relative to the external surface of the cast component 60.
  • the projections 32 attached to the core 10 ensure that the cooling holes are directly located relative to the internal surface.
  • the projections 32 and dimples or recesses 36 may also inhibit movement of the core 10 during the method of forming described herein. Although four projections 32 are shown, this is merely exemplaiy as any number of cooling holes may be formed depending on the flow rate and cooling capacity needed.
  • the core 10 is in fluid communication with a port 19 into the die 34.
  • the port 19 allows fluid flow of wax 21 into the die 34 and about the core 10 therein.
  • the wax 21 is also shown in the die 34 surrounding the core 10. The wax 21 forms about the projections 32 and within the dimples 36.
  • FIG. 4 a schematic view of the wax pattern 38 is shown with the wax projections 39 formed by way of the projections 32 of the core 10, inside the wax pattern 38, and the dimples 36 located within the die.
  • a plurality of these wax patterns 38 are constructed and formed into a tree 46 as depicted in FIG. 5.
  • the tree 46 of exemplary embodiment includes a plurality of wax patterns 38. According to the exemplary embodiment, six wax patterns 38 are shown defining the exemplary tree 46. However, various numbers may be utilized during this manufacturing step,
  • the wax pattern tree 46 is wetted by an agent 48.
  • the agent maybe zireonia slurry however, alternate wetting agents may be utilized.
  • the tree 46 is shown being dipped into a bath. However, as alternative means of manufacturing, a spray may be utilized to thoroughly wet the patterns 38 of the tree 46.
  • the wetting agent 48 provides a means of adhering ceramic material to the wax patterns 38.
  • a stuccoing process occurs wherein the wetted wax patterns 38 are coated with ceramic 29.
  • the ceramic may include a zirconium, silicon or aluminum based sand material. However, alternate ceramics may be utilized. This process of wetting and stuccoing may occur one or more times until a desired thickness of the stucco coating is obtained and the ceramic is sufficiently hardened for further steps in the process.
  • the ceramic mold 40 has projections formed along the inner surface corresponding to the wax porjections of the wax patterns 38. At this point the mold is dried in preparation for further processing.
  • the tree 46 with the molds 40 is positioned within an oven 27 for de- waxing.
  • the temperature of the tree 46 is raised so that the wax 21 of the pattern 38 melts and is released from the interior of the tree 46.
  • steam may be utilized to melt wax internally from the molds.
  • the de- waxing may continue in a bum out process depicted in FIG. 9 wherein residue wax is removed and the stucco shells 40 are fired for curing. This may occur in the oven depicted in FIG. 8 or may occur in a separate structure depending on the manufacturing process.
  • a molten metal is poured into the shell 40.
  • the molten metal 42 fills the interior of the ceramic shell 40 and hardens to form a metallic part inside the shell 40.
  • the molten metal which is cast in this manner may be any of the standard alloys.
  • aluminum, nickel or iron based alloys can be used.
  • the resulting cast structures are varied, for example equiaxed grain structures, directionally solidified grain struclisres and single crystal grain structures could be cast using this process.
  • the stucco or ceramic shell 40 is removed. This is usually accomplished by mechanically removing the shell although other processes may be used.
  • the trees 46 may be shaken to break the shell 40 revealing the cast component 60.
  • the ceramic shell material 40 is shown falling and broken in pieces from the tree 46.
  • the cast components 60 are removed fro the tree 46. This may occur in a variety of processes.
  • a cutting tool is utilized to remove the components 60 from the tree 46.
  • Various types of cutting tools may be utilized.
  • the projections 32 defined in the initial steps of FIGS. 2-4 are still present, in the cast component 60. With the ceramic removed from the cast component 60, the metallic material encloses the projections which were formed.
  • the casting is exposed to a leaching solution to remove the ceramic core.
  • the leaching solution is usually a caustic liquid and may also utilize pressure and temperature to accelerate the leaching.
  • Fi(l 14 after the part 60 is removed from the tree 46, the part must be finished. This finishing process may occur in a variety of manners. According to the exemplary embodiment, one non-limiting example of the finishing process includes a surface grind. In the finishing process, the metal of the cast component 60 is removed at the location of the projections previously described. In the example depicted, the projections 62 are ground to reveal the cooling holes 64.
  • These core cooling holes provide airflow communication from the inside of the component 60 to the exterior with the metallic tips of the projections 62 removed.
  • the cooling holes are located properly during the casting process of the part and are not subject to problems associated with machining subsequently, such as mislocation of the hole or the like.
  • FIG. 15 the component 60 consists of the cast metal projections 62 and the void or cooling hole 64 where the core 10 was before leaching.
  • Figure 15 shows the article before finishing.
  • Figure 16 shows the component 60 after finishing.
  • FIG. 16 shows the metallic part with an opening 64 defining the cooling hole and providing access to the interior of the cast component 60.
  • the interior of the part 60 is shown below the cross-hatch and the exterior of the part is above the cross-hatch.
  • the method 100 first requires positioning a core in the die at step 102.
  • wax is injected into the die to form a pattern at step 104.
  • the wax pattern is coated with ceramic at step 106.
  • a subsequent de-waxing occurs at step 108 wherein the wax is removed from the interior of the ceramic shell.
  • the component is cast by pouring molten metal into the ceramic shell and allowing the metal to solidify.
  • the ceramic shell is removed at step 1 12. This may occur in a variety of means as previously described.
  • the component is removed from the tree at step 1 14.
  • the core 10 is leached out of the cast component at step 1 16.
  • the component is finished at step 1 18to reveal the one or more cooling apertures.
  • inventive embodiments are presented by way of example only and that, within the scope of the appended claims and equivalents thereto, inventive embodiments may be practiced otherwise than as specifically described and claimed.
  • inventive embodiments of the present disclosure are directed to each individual feature, system, article, material, kit, and/or method described herein.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Abstract

L'invention concerne un procédé permettant de former des orifices de refroidissement coulés dans un composant d'un moteur d'aéronef. Le procédé consiste à utiliser un mandrin pourvu de projections et un poinçon pourvu d'alvéoles ou d'évidements pour former les orifices de refroidissement coulés, et une étape de finissage pour révéler les orifices de refroidissement coulés.
PCT/US2013/076906 2013-01-18 2013-12-20 Procédé permettant de former des orifices de refroidissement coulés dans un composant d'aéronef WO2014113184A1 (fr)

Applications Claiming Priority (2)

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US201313744740A 2013-01-18 2013-01-18
US13/744,740 2013-01-18

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WO2014113184A1 true WO2014113184A1 (fr) 2014-07-24

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Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10022790B2 (en) * 2014-06-18 2018-07-17 Siemens Aktiengesellschaft Turbine airfoil cooling system with leading edge impingement cooling system turbine blade investment casting using film hole protrusions for integral wall thickness control
US10408079B2 (en) 2015-02-18 2019-09-10 Siemens Aktiengesellschaft Forming cooling passages in thermal barrier coated, combustion turbine superalloy components
US10927705B2 (en) 2018-08-17 2021-02-23 Raytheon Technologies Corporation Method for forming cooling holes having separate complex and simple geometry sections
CN113600755A (zh) * 2021-08-31 2021-11-05 中国航发沈阳黎明航空发动机有限责任公司 一种带测温孔多联体叶片的铸造方法
FR3137316A1 (fr) * 2022-06-29 2024-01-05 Safran Aircraft Engines Noyau céramique pour aube de turbine creuse à perçages externes

Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5291654A (en) * 1993-03-29 1994-03-08 United Technologies Corporation Method for producing hollow investment castings
US5853044A (en) * 1996-04-24 1998-12-29 Pcc Airfoils, Inc. Method of casting an article
US20060162893A1 (en) * 2004-12-27 2006-07-27 Thomas Beck Method for the production of a casting mold
US7172012B1 (en) * 2004-07-14 2007-02-06 United Technologies Corporation Investment casting
US20070044936A1 (en) * 2005-09-01 2007-03-01 United Technologies Corporation Cooled turbine airfoils and methods of manufacture

Patent Citations (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US5291654A (en) * 1993-03-29 1994-03-08 United Technologies Corporation Method for producing hollow investment castings
US5853044A (en) * 1996-04-24 1998-12-29 Pcc Airfoils, Inc. Method of casting an article
US7172012B1 (en) * 2004-07-14 2007-02-06 United Technologies Corporation Investment casting
US20060162893A1 (en) * 2004-12-27 2006-07-27 Thomas Beck Method for the production of a casting mold
US20070044936A1 (en) * 2005-09-01 2007-03-01 United Technologies Corporation Cooled turbine airfoils and methods of manufacture

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10022790B2 (en) * 2014-06-18 2018-07-17 Siemens Aktiengesellschaft Turbine airfoil cooling system with leading edge impingement cooling system turbine blade investment casting using film hole protrusions for integral wall thickness control
US10408079B2 (en) 2015-02-18 2019-09-10 Siemens Aktiengesellschaft Forming cooling passages in thermal barrier coated, combustion turbine superalloy components
US10927705B2 (en) 2018-08-17 2021-02-23 Raytheon Technologies Corporation Method for forming cooling holes having separate complex and simple geometry sections
CN113600755A (zh) * 2021-08-31 2021-11-05 中国航发沈阳黎明航空发动机有限责任公司 一种带测温孔多联体叶片的铸造方法
FR3137316A1 (fr) * 2022-06-29 2024-01-05 Safran Aircraft Engines Noyau céramique pour aube de turbine creuse à perçages externes

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