WO2019099000A1 - Method of repairing gamma prime strengthened superalloys - Google Patents

Method of repairing gamma prime strengthened superalloys Download PDF

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Publication number
WO2019099000A1
WO2019099000A1 PCT/US2017/061722 US2017061722W WO2019099000A1 WO 2019099000 A1 WO2019099000 A1 WO 2019099000A1 US 2017061722 W US2017061722 W US 2017061722W WO 2019099000 A1 WO2019099000 A1 WO 2019099000A1
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WIPO (PCT)
Prior art keywords
filler material
defective portion
gas turbine
turbine engine
engine component
Prior art date
Application number
PCT/US2017/061722
Other languages
French (fr)
Inventor
Gerald J. Bruck
Original Assignee
Siemens Energy, Inc.
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 Siemens Energy, Inc. filed Critical Siemens Energy, Inc.
Priority to PCT/US2017/061722 priority Critical patent/WO2019099000A1/en
Publication of WO2019099000A1 publication Critical patent/WO2019099000A1/en

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Classifications

    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B23—MACHINE TOOLS; METAL-WORKING NOT OTHERWISE PROVIDED FOR
    • B23P—METAL-WORKING NOT OTHERWISE PROVIDED FOR; COMBINED OPERATIONS; UNIVERSAL MACHINE TOOLS
    • B23P6/00—Restoring or reconditioning objects
    • B23P6/002—Repairing turbine components, e.g. moving or stationary blades, rotors
    • B23P6/007—Repairing turbine components, e.g. moving or stationary blades, rotors using only additive methods, e.g. build-up welding
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B22—CASTING; POWDER METALLURGY
    • B22F—WORKING METALLIC POWDER; MANUFACTURE OF ARTICLES FROM METALLIC POWDER; MAKING METALLIC POWDER; APPARATUS OR DEVICES SPECIALLY ADAPTED FOR METALLIC POWDER
    • B22F10/00—Additive manufacturing of workpieces or articles from metallic powder
    • B22F10/20—Direct sintering or melting
    • B22F10/28—Powder bed fusion, e.g. selective laser melting [SLM] or electron beam melting [EBM]
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B22—CASTING; POWDER METALLURGY
    • B22F—WORKING METALLIC POWDER; MANUFACTURE OF ARTICLES FROM METALLIC POWDER; MAKING METALLIC POWDER; APPARATUS OR DEVICES SPECIALLY ADAPTED FOR METALLIC POWDER
    • B22F10/00—Additive manufacturing of workpieces or articles from metallic powder
    • B22F10/60—Treatment of workpieces or articles after build-up
    • B22F10/64—Treatment of workpieces or articles after build-up by thermal means
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B22—CASTING; POWDER METALLURGY
    • B22F—WORKING METALLIC POWDER; MANUFACTURE OF ARTICLES FROM METALLIC POWDER; MAKING METALLIC POWDER; APPARATUS OR DEVICES SPECIALLY ADAPTED FOR METALLIC POWDER
    • B22F3/00—Manufacture of workpieces or articles from metallic powder characterised by the manner of compacting or sintering; Apparatus specially adapted therefor ; Presses and furnaces
    • B22F3/12—Both compacting and sintering
    • B22F3/14—Both compacting and sintering simultaneously
    • B22F3/15—Hot isostatic pressing
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B22—CASTING; POWDER METALLURGY
    • B22F—WORKING METALLIC POWDER; MANUFACTURE OF ARTICLES FROM METALLIC POWDER; MAKING METALLIC POWDER; APPARATUS OR DEVICES SPECIALLY ADAPTED FOR METALLIC POWDER
    • B22F5/00—Manufacture of workpieces or articles from metallic powder characterised by the special shape of the product
    • B22F5/04—Manufacture of workpieces or articles from metallic powder characterised by the special shape of the product of turbine blades
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B22—CASTING; POWDER METALLURGY
    • B22F—WORKING METALLIC POWDER; MANUFACTURE OF ARTICLES FROM METALLIC POWDER; MAKING METALLIC POWDER; APPARATUS OR DEVICES SPECIALLY ADAPTED FOR METALLIC POWDER
    • B22F7/00—Manufacture of composite layers, workpieces, or articles, comprising metallic powder, by sintering the powder, with or without compacting wherein at least one part is obtained by sintering or compression
    • B22F7/06—Manufacture of composite layers, workpieces, or articles, comprising metallic powder, by sintering the powder, with or without compacting wherein at least one part is obtained by sintering or compression of composite workpieces or articles from parts, e.g. to form tipped tools
    • B22F7/062—Manufacture of composite layers, workpieces, or articles, comprising metallic powder, by sintering the powder, with or without compacting wherein at least one part is obtained by sintering or compression of composite workpieces or articles from parts, e.g. to form tipped tools involving the connection or repairing of preformed parts
    • 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
    • F01D5/005—Repairing methods or devices
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B22—CASTING; POWDER METALLURGY
    • B22F—WORKING METALLIC POWDER; MANUFACTURE OF ARTICLES FROM METALLIC POWDER; MAKING METALLIC POWDER; APPARATUS OR DEVICES SPECIALLY ADAPTED FOR METALLIC POWDER
    • B22F7/00—Manufacture of composite layers, workpieces, or articles, comprising metallic powder, by sintering the powder, with or without compacting wherein at least one part is obtained by sintering or compression
    • B22F7/06—Manufacture of composite layers, workpieces, or articles, comprising metallic powder, by sintering the powder, with or without compacting wherein at least one part is obtained by sintering or compression of composite workpieces or articles from parts, e.g. to form tipped tools
    • B22F7/062—Manufacture of composite layers, workpieces, or articles, comprising metallic powder, by sintering the powder, with or without compacting wherein at least one part is obtained by sintering or compression of composite workpieces or articles from parts, e.g. to form tipped tools involving the connection or repairing of preformed parts
    • B22F2007/068—Manufacture of composite layers, workpieces, or articles, comprising metallic powder, by sintering the powder, with or without compacting wherein at least one part is obtained by sintering or compression of composite workpieces or articles from parts, e.g. to form tipped tools involving the connection or repairing of preformed parts repairing articles
    • B—PERFORMING OPERATIONS; TRANSPORTING
    • B22—CASTING; POWDER METALLURGY
    • B22F—WORKING METALLIC POWDER; MANUFACTURE OF ARTICLES FROM METALLIC POWDER; MAKING METALLIC POWDER; APPARATUS OR DEVICES SPECIALLY ADAPTED FOR METALLIC POWDER
    • B22F5/00—Manufacture of workpieces or articles from metallic powder characterised by the special shape of the product
    • B22F5/009—Manufacture of workpieces or articles from metallic powder characterised by the special shape of the product of turbine components other than turbine blades
    • C—CHEMISTRY; METALLURGY
    • C22—METALLURGY; FERROUS OR NON-FERROUS ALLOYS; TREATMENT OF ALLOYS OR NON-FERROUS METALS
    • C22C—ALLOYS
    • C22C1/00—Making non-ferrous alloys
    • C22C1/04—Making non-ferrous alloys by powder metallurgy
    • C22C1/0433—Nickel- or cobalt-based alloys
    • C—CHEMISTRY; METALLURGY
    • C22—METALLURGY; FERROUS OR NON-FERROUS ALLOYS; TREATMENT OF ALLOYS OR NON-FERROUS METALS
    • C22C—ALLOYS
    • C22C32/00—Non-ferrous alloys containing at least 5% by weight but less than 50% by weight of oxides, carbides, borides, nitrides, silicides or other metal compounds, e.g. oxynitrides, sulfides, whether added as such or formed in situ
    • C22C32/001—Non-ferrous alloys containing at least 5% by weight but less than 50% by weight of oxides, carbides, borides, nitrides, silicides or other metal compounds, e.g. oxynitrides, sulfides, whether added as such or formed in situ with only oxides
    • C22C32/0015—Non-ferrous alloys containing at least 5% by weight but less than 50% by weight of oxides, carbides, borides, nitrides, silicides or other metal compounds, e.g. oxynitrides, sulfides, whether added as such or formed in situ with only oxides with only single oxides as main non-metallic constituents
    • C22C32/0026—Matrix based on Ni, Co, Cr or alloys thereof
    • C—CHEMISTRY; METALLURGY
    • C22—METALLURGY; FERROUS OR NON-FERROUS ALLOYS; TREATMENT OF ALLOYS OR NON-FERROUS METALS
    • C22C—ALLOYS
    • C22C33/00—Making ferrous alloys
    • C22C33/02—Making ferrous alloys by powder metallurgy
    • C22C33/0257—Making ferrous alloys by powder metallurgy characterised by the range of the alloying elements
    • C22C33/0278—Making ferrous alloys by powder metallurgy characterised by the range of the alloying elements with at least one alloying element having a minimum content above 5%
    • C22C33/0285—Making ferrous alloys by powder metallurgy characterised by the range of the alloying elements with at least one alloying element having a minimum content above 5% with Cr, Co, or Ni having a minimum content higher than 5%
    • 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
    • F05D2230/22—Manufacture essentially without removing material by sintering
    • 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
    • F05D2300/00—Materials; Properties thereof
    • F05D2300/10—Metals, alloys or intermetallic compounds
    • F05D2300/17—Alloys
    • F05D2300/175—Superalloys
    • 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
    • Y02P—CLIMATE CHANGE MITIGATION TECHNOLOGIES IN THE PRODUCTION OR PROCESSING OF GOODS
    • Y02P10/00—Technologies related to metal processing
    • Y02P10/25—Process efficiency

Definitions

  • Disclosed embodiments are generally related to repairing of gas turbine engine components and in particular to the repairing of gas turbine engine components made of gamma prime strengthened superalloys.
  • Cobalt based superalloys may perform better than nickel based superalloys due to having a higher melting point, a higher stress rupture strength at elevated temperatures, better hot corrosion resistance due to higher Cr content, easier to weld, and better fatigue resistance.
  • the cobalt based superalloys can in some ways perform worse than nickel based superalloys due to lower strength, lower ductility, lower fracture toughness and limited potential to improve their properties.
  • cobalt based superalloys can be a result of limited strengthening mechanisms.
  • Most cobalt based superalloys depend on solid solution strengthening from elements such as Cr, Ta, W, Nb, Mo and from carbide precipitation hardening.
  • Nickel based superalloys achieve superior strength to high temperatures from gamma prime [Ni 3 (Al, Ti)] and gamma double prime [NTCb] precipitations.
  • gamma prime strengthened superalloys have been more recently disclosed.
  • a nickel and chromium free cobalt-based superalloy providing strength from gamma prime has been developed with strengthening precipitate defined as Co3(Al,W). Tantalum is added to stabilize the gamma prime.
  • This superalloy has high temperature strength and oxidation resistance and is expected to have gamma prime that is more stable than the gamma prime in nickel-base superalloys.
  • dispersion stability may be a positive attribute for superalloy performance, it proves to be a challenge for repairs.
  • a popular technique for repair of nickel-based superalloys is to over age heat treat the material to grow gamma prime size, reduce hardness and improve the ability to repair it. With advanced stabilized gamma prime cobalt-based superalloys, such an over age treatment is not as effective.
  • aspects of the present disclosure relate to repairing gamma prime strengthened superalloys.
  • An aspect of the present disclosure may be a method for repairing a gas turbine engine component.
  • the method may comprise removing a defective portion of the gas turbine engine component, wherein the defective portion is made of a gamma prime strengthened superalloy material; depositing filler material in a cavity left by the removal of the defective portion; sintering the filler material with selective laser sintering; repeating depositing the filler material and selective layer sintering the filler material until the cavity is filled or overfilled with the filler material ; sealing a last layer of the filler material by forming a fully dense skin over the filler material with selective laser melting; and hot isotactic pressing the filler material.
  • Another aspect of the present disclosure may be a method for repairing a gas turbine engine component.
  • the method may comprise removing a defective portion of the gas turbine engine component, wherein the defective portion is made of a gamma prime strengthened superalloy material; depositing filler material in a cavity left by the removal of the defective portion; sintering the filler material with selective laser sintering; depositing the filler material and selective layer sintering the filler material until the cavity is filled or overfilled with the filler material; sealing a last layer of the filler material by melting and fully densifying a ductile alloy skin over the last layer of the filler material; and hot isotactic pressing the filler material.
  • Yet another aspect of the present invention may be a method for repairing a gas turbine engine component.
  • the method may comprise removing a defective portion of the gas turbine engine component, wherein the defective portion is made of a gamma prime strengthened cobalt based superalloy material; depositing filler material in a cavity left by the removal of the defective portion; sintering the filler material with selective laser sintering; depositing the filler material and selective layer sintering the filler material until the cavity is filled or overfilled with the filler material; sealing a last layer of the filler material by forming a fully dense skin over the filler material with selective laser melting; and hot isotactic pressing the filler material.
  • Fig. 1 is a flow chart setting forth the method for repairing a gas turbine engine component.
  • Fig. 2 shows a gas turbine engine component having a defective portion.
  • FIG. 3 shows the gas turbine engine component with the defective portion removed.
  • Fig. 4 shows a gas turbine engine component having filler material placed and sintered in the area where the defective portion was removed.
  • Fig. 5 shows a lOskin formed on the filler material via the usage of selective laser melting.
  • Fig. 6 shows the gas turbine engine component after the application of hot isotactic pressing.
  • Fig. 7 is a flow chart setting forth an alternative method for repairing a gas turbine engine.
  • Fig. 8 shows a gas turbine engine component having a defective portion.
  • Fig. 9 shows the gas turbine engine component with the defective portion removed.
  • Fig. 10 shows a gas turbine engine component having filler material placed and sintered in the area where the defective portion was removed.
  • Fig. 11 shows a ductile alloy skin formed on the filler material via the usage of selective laser melting.
  • Fig. 12 shows the gas turbine engine component after the application of hot isotactic pressing.
  • Fig. 13 shows the gas turbine engine component after the removal of the ductile alloy skin.
  • Fig. 1 is a flow chart setting forth the method for repairing a gas turbine engine. Discussion of the method for repairing a gas turbine engine will be discussed with reference to Figs. 2-6, wherein a gas turbine engine component 10 is shown that is defective in some manner.
  • the gas turbine engine component 10 may be part of a combustor, a stator, a vane, a transition piece, etc.
  • gas turbine engine component 10 While a gas turbine engine component 10 is discussed herein it should be understood that this method may be used for other types of components for other types of machines, provided the same types of materials and conditions are involved.
  • the gas turbine engine component 10 is made of a gamma prime strengthened cobalt-based superalloy. However, other types of alloys may be used as well, such as nickel, nickel-cobalt, iron, iron-nickel-based and oxide dispersion strengthened superalloys.
  • Fig. 2 shows the gas turbine engine component 10 having a defective portion 12.
  • the defective portion 12 may be a pre-existing flaw that exists in the gas turbine engine component 10, or may be the result of damage (e.g. foreign object impact, fatigue, corrosion, creep) that occurred during the operation of the gas turbine engine.
  • the defective portion 12 may be removed from the gas turbine engine component 10. This may be achieved via excavation of the portion using grinding, ball mill machining, electro discharge machining, air arc gouging and other similar commercial methods of material removal.
  • FIG. 3 shows the gas turbine engine component 10 with the defective portion 12 removed. Shown is the cavity 14 that previously held the defective portion 12
  • step 104 filler material 16 is placed in the cavity 14.
  • Fig. 4 shows a gas turbine engine component 10 with the filler material 16 placed in the cavity 14 where the defective portion 12 was removed.
  • the filler material 16 may be the same material as the gas turbine engine component 10. However, as discussed below, the filler material 16 may be made of a different superalloy.
  • step 106 the filler material 16 undergoes selective laser sintering.
  • the selective laser sintering fills and preferably overfills the cavity 14 with the filler material 16.
  • Steps 104 and 106 preferably involve placing and sintering material in relatively thin, e.g. 20 to 100 micron thick, layers and repeating such layered placement and sintering to build up and fill the cavity 14.
  • the cavity 14 is overfilled so that hot isostatic pressing collapses the overfilled cavity 14 and closes spaces between sintered particles of the filler material 16. This results in a final consolidation and complete filling of the then-repaired volume.
  • the overfilling needs to be sufficient in size to compensate for the volume of spaces left between sintered particles of the filler material 16.
  • step 108 the last layer 17 of the filler material 16 undergoes selective laser melting.
  • Fig. 5 shows a fully dense skin 18 formed on the filler material 16 via the usage of selective laser melting. Densification is accomplished by melting and solidification of the last layer 17. A“fully dense skin” has from zero to less than five percent porosity.
  • the selective laser melting of the last layer 17 of the filler material 16 forms a skin 18.
  • the skin 18 is preferably porositiy-free. Such absence of porosity is important for the next step of consolidation by hot isostatic pressing (HIP). That is, (as is known in metal casting industry practice) surface pores are not generally closed by the HIP process, a solid skin ensures pores at and near to (e.g. within a layer thickness of) the surface get closed.
  • HIP hot isostatic pressing
  • connection point 19 between the skin 18 and the remainder of the gas turbine engine component should not adversely affect the gas turbine engine component 10.
  • the skin 18 and the adjoining melted area of the gas turbine engine component 10 should be no thicker than 0.5 mm, preferably no thicker than 0.1 mm and most advantageously within the range of 25 to 100 microns.
  • the connection 19 is of sufficient strength to survive the process that it undergoes in step 110.
  • the thickness of the skin 18 is of no consequence so long as the process discussed below causes the skin 18 to collapse on the remainder of the material of the gas turbine engine component 10.
  • step 110 the filler material 16 of the gas turbine engine component 10 goes through hot isotactic pressing. This is to fully compress the filler material 16 within the cavity 14.
  • Fig. 6 shows the gas turbine engine component 10 after the application of hot isotactic pressing. After the hot isotactic pressing that occurs in step 110, the repaired gas turbine engine component can undergo machining and inspection prior to being returned to the gas turbine engine.
  • FIG. 7 a flow chart setting forth an alternative method for repairing a gas turbine engine component is shown.
  • the flow chart and method set forth in Fig. 7 is similar to the method set forth above. The differences address a potential issue of the embodiment found in Figure 5. This is accomplished in that the connection 19 could be too fragile if the filler material 16 in this situation was that same as that of the remainder of the gas turbine engine component 10. So in this case the filler material 16 is chosen to be matching or a superalloy material, but the final layer 17 ( Figure 11) is chosen to be a ductile alloy. This can occur in situations where the gas turbine engine component 10 is made of a cast nickel based superalloy such as Inconel 939, Inconel 738 Rene 80, CM 247 or case cobalt based alloy such as earlier described.
  • a cast nickel based superalloy such as Inconel 939, Inconel 738 Rene 80, CM 247 or case cobalt based alloy such as earlier described.
  • Fig. 8 shows the gas turbine engine component 10 having a defective portion 12.
  • the defective portion 12 may a pre-existing flaw that exists in the gas turbine engine component 10, or may be the result of damage that occurred during the operation of the gas turbine engine.
  • the defective portion 12 may be removed from the gas turbine engine component 10. This may be achieved via excavation of the portion using grinding, ball mill machining, electro discharge machining, air arc gouging and other similar commercial methods of material removal.
  • Fig. 9 shows the gas turbine engine component 10 with the defective portion 12 removed. Shown is the cavity 14 that previously held the defective portion 12 of the gas turbine engine component 10.
  • step 204 filler material 16 is placed in the cavity 14.
  • Fig. 10 shows a gas turbine engine component 10 having the filler material 16 placed in the cavity 14 where the defective portion 12 was removed.
  • the filler material 16 may be a matching or alternate superalloy in this embodiment.
  • the last layer is deposited of a sacrificial ductile alloy that is readily welded to the substrate.
  • the ductile alloy may be Inconel alloy 600, 625, 617, 82 or Haynes alloy X, 282 , a stainless steel such as 304, 316, 347 or a ductile cobalt based alloy such as Haynes 188, L-605, or Mar-M 918.
  • the filler material 16 undergoes selective laser sintering.
  • the selective laser sintering fills and preferably overfills the cavity 14 with filler material 16.
  • Steps 204 and 206 preferably involve placing and sintering material in relatively thin, e.g. 20 to 100 micron thick, layers and repeating such layered placement and sintering to build up and fill the cavity 14.
  • step 208 the last layer 17 of the final ductile layer of filler material 16 undergoes selective laser melting and full densification.
  • Fig. 11 shows a skin 18 formed on the filler material 16 via the usage of selective laser melting of the ductile alloy.
  • the selective laser melting of the last layer 17 of the filler material 16 forms a skin 18.
  • the skin 18 is preferably porositiy-free.
  • connection point 19 between the skin 18 and the remainder of the gas turbine engine component should not adversely affect the gas turbine engine component 10.
  • the skin 18 and the adjoining melted area of the gas turbine engine 10 should be no thicker than 0.5 mm, preferably no thicker than 0.1 mm and most advantageously within the range of 25 to 100 microns.
  • the connection 19 is of sufficient strength to survive the process that it undergoes in step 210.
  • the filler material 16 of the gas turbine engine component 10 goes through hot isotactic pressing. This is to fully compress the filler material 16 within the cavity 14.
  • Fig. 12 shows the gas turbine engine component 10 after the application of hot isotactic pressing.
  • step 212 the skin 18 that is formed from the ductile alloy may be removed.
  • the repaired gas turbine engine component 10 can undergo machining and inspection prior to being returned to the gas turbine engine.

Abstract

A gas turbine engine component (10) is repaired via the removal of the defective portion (12) and then the filling in of the resulting cavity (14) with filler material (16). The filler material (16) fills in the cavity (14) via selective laser sintering. A skin (18) is formed on the filler material (16) via selective laser melting. The gas turbine engine component then undergoes hot isotactic pressing. This method is used with gamma prime strengthened superalloy materials.

Description

METHOD OF REPAIRING GAMMA PRIME STRENGTHENED SUPERALLOYS
BACKGROUND
[0001] 1. Field
[0002] Disclosed embodiments are generally related to repairing of gas turbine engine components and in particular to the repairing of gas turbine engine components made of gamma prime strengthened superalloys.
[0003] 2. Description of the Related Art
[0004] Cobalt based superalloys may perform better than nickel based superalloys due to having a higher melting point, a higher stress rupture strength at elevated temperatures, better hot corrosion resistance due to higher Cr content, easier to weld, and better fatigue resistance. However, the cobalt based superalloys can in some ways perform worse than nickel based superalloys due to lower strength, lower ductility, lower fracture toughness and limited potential to improve their properties.
[0005] The lower strength of cobalt based superalloys can be a result of limited strengthening mechanisms. Most cobalt based superalloys depend on solid solution strengthening from elements such as Cr, Ta, W, Nb, Mo and from carbide precipitation hardening. Nickel based superalloys achieve superior strength to high temperatures from gamma prime [Ni3(Al, Ti)] and gamma double prime [NTCb] precipitations.
[0006] Early attempts at gamma prime strengthened cobalt based superalloys have shown unfavourable properties. One example is a cobalt-based superalloy formed with nickel, chromium, titanium and aluminum. The gamma prime precipitation that forms is (Co,Ni)3Ti. Unfortunately, the precipitate has a hexagonal plate-like structure and provides inferior high temperature properties thereby limiting the use of the superalloy.
[0007] Other gamma prime strengthened superalloys have been more recently disclosed. A nickel and chromium free cobalt-based superalloy providing strength from gamma prime has been developed with strengthening precipitate defined as Co3(Al,W). Tantalum is added to stabilize the gamma prime. This superalloy has high temperature strength and oxidation resistance and is expected to have gamma prime that is more stable than the gamma prime in nickel-base superalloys. [0008] While such dispersion stability may be a positive attribute for superalloy performance, it proves to be a challenge for repairs. A popular technique for repair of nickel-based superalloys is to over age heat treat the material to grow gamma prime size, reduce hardness and improve the ability to repair it. With advanced stabilized gamma prime cobalt-based superalloys, such an over age treatment is not as effective.
[0009] Another limitation with the cobalt-based gamma prime strengthened superalloy relates to its near constant low level of ductility. Even up to 900 °C the elongation of the gamma prime strengthened superalloy is only about 3 percent. While such elongation may be of little consequence during service where strength is of overriding importance, it provides a challenge for repairs. A popular technique for repair of nickel-based superalloys is to heat the component to elevated temperatures to achieve improved ductility and then to repair weld at the elevated temperatures. This process is widely referred to as SWET (superalloy welding at elevated temperature). However for cobalt-based gamma prime strengthened superalloys, welding at elevated temperature may be ineffective.
[0010] Therefore, a method for effectively repairing cobalt-based gamma prime strengthened superalloys is needed.
SUMMARY
[0011] Briefly described, aspects of the present disclosure relate to repairing gamma prime strengthened superalloys.
[0012] An aspect of the present disclosure may be a method for repairing a gas turbine engine component. The method may comprise removing a defective portion of the gas turbine engine component, wherein the defective portion is made of a gamma prime strengthened superalloy material; depositing filler material in a cavity left by the removal of the defective portion; sintering the filler material with selective laser sintering; repeating depositing the filler material and selective layer sintering the filler material until the cavity is filled or overfilled with the filler material ; sealing a last layer of the filler material by forming a fully dense skin over the filler material with selective laser melting; and hot isotactic pressing the filler material.
[0013] Another aspect of the present disclosure may be a method for repairing a gas turbine engine component. The method may comprise removing a defective portion of the gas turbine engine component, wherein the defective portion is made of a gamma prime strengthened superalloy material; depositing filler material in a cavity left by the removal of the defective portion; sintering the filler material with selective laser sintering; depositing the filler material and selective layer sintering the filler material until the cavity is filled or overfilled with the filler material; sealing a last layer of the filler material by melting and fully densifying a ductile alloy skin over the last layer of the filler material; and hot isotactic pressing the filler material.
[0014] Yet another aspect of the present invention may be a method for repairing a gas turbine engine component. The method may comprise removing a defective portion of the gas turbine engine component, wherein the defective portion is made of a gamma prime strengthened cobalt based superalloy material; depositing filler material in a cavity left by the removal of the defective portion; sintering the filler material with selective laser sintering; depositing the filler material and selective layer sintering the filler material until the cavity is filled or overfilled with the filler material; sealing a last layer of the filler material by forming a fully dense skin over the filler material with selective laser melting; and hot isotactic pressing the filler material.
BRIEF DESCRIPTION OF THE DRAWINGS
[0015] Fig. 1 is a flow chart setting forth the method for repairing a gas turbine engine component.
[0016] Fig. 2 shows a gas turbine engine component having a defective portion.
[0017] Fig. 3 shows the gas turbine engine component with the defective portion removed.
[0018] Fig. 4 shows a gas turbine engine component having filler material placed and sintered in the area where the defective portion was removed.
[0019] Fig. 5 shows a lOskin formed on the filler material via the usage of selective laser melting. [0020] Fig. 6 shows the gas turbine engine component after the application of hot isotactic pressing.
[0021] Fig. 7 is a flow chart setting forth an alternative method for repairing a gas turbine engine.
[0022] Fig. 8 shows a gas turbine engine component having a defective portion.
[0023] Fig. 9 shows the gas turbine engine component with the defective portion removed.
[0024] Fig. 10 shows a gas turbine engine component having filler material placed and sintered in the area where the defective portion was removed.
[0025] Fig. 11 shows a ductile alloy skin formed on the filler material via the usage of selective laser melting.
[0026] Fig. 12 shows the gas turbine engine component after the application of hot isotactic pressing.
[0027] Fig. 13 shows the gas turbine engine component after the removal of the ductile alloy skin.
DETAILED DESCRIPTION
[0028] To facilitate an understanding of embodiments, principles, and features of the present disclosure, they are disclosed hereinafter with reference to implementation in illustrative embodiments. Embodiments of the present disclosure, however, are not limited to use in the described systems or methods and may be utilized in other systems and methods as will be understood by those skilled in the art.
[0029] The various embodiments are intended to be illustrative and not restrictive. Many suitable mixtures that would perform the same or a similar function as the additive manufacturing powders described herein are intended to be embraced within the scope of embodiments of the present disclosure.
[0030] This disclosure provides an improved method to repair advanced, gamma prime strengthened superalloy components, in particular gamma prime strengthened cobalt-based superalloy components. It avoids ineffective methods including over age pre-weld heat treatment and repair welding at elevated temperatures. [0031] Fig. 1 is a flow chart setting forth the method for repairing a gas turbine engine. Discussion of the method for repairing a gas turbine engine will be discussed with reference to Figs. 2-6, wherein a gas turbine engine component 10 is shown that is defective in some manner. The gas turbine engine component 10 may be part of a combustor, a stator, a vane, a transition piece, etc. While a gas turbine engine component 10 is discussed herein it should be understood that this method may be used for other types of components for other types of machines, provided the same types of materials and conditions are involved. For the purpose of discussion the gas turbine engine component 10 is made of a gamma prime strengthened cobalt-based superalloy. However, other types of alloys may be used as well, such as nickel, nickel-cobalt, iron, iron-nickel-based and oxide dispersion strengthened superalloys.
[0032] Fig. 2 shows the gas turbine engine component 10 having a defective portion 12. The defective portion 12 may be a pre-existing flaw that exists in the gas turbine engine component 10, or may be the result of damage (e.g. foreign object impact, fatigue, corrosion, creep) that occurred during the operation of the gas turbine engine. Upon discovery of the defective portion, in step 102 the defective portion 12 may be removed from the gas turbine engine component 10. This may be achieved via excavation of the portion using grinding, ball mill machining, electro discharge machining, air arc gouging and other similar commercial methods of material removal.
[0033] Fig. 3 shows the gas turbine engine component 10 with the defective portion 12 removed. Shown is the cavity 14 that previously held the defective portion 12
[0034] In step 104, filler material 16 is placed in the cavity 14. Fig. 4 shows a gas turbine engine component 10 with the filler material 16 placed in the cavity 14 where the defective portion 12 was removed. The filler material 16 may be the same material as the gas turbine engine component 10. However, as discussed below, the filler material 16 may be made of a different superalloy.
[0035] In step 106, the filler material 16 undergoes selective laser sintering. The selective laser sintering fills and preferably overfills the cavity 14 with the filler material 16. Steps 104 and 106 preferably involve placing and sintering material in relatively thin, e.g. 20 to 100 micron thick, layers and repeating such layered placement and sintering to build up and fill the cavity 14. As sintering leaves spaces between particles of the filler material 16, preferably the cavity 14 is overfilled so that hot isostatic pressing collapses the overfilled cavity 14 and closes spaces between sintered particles of the filler material 16. This results in a final consolidation and complete filling of the then-repaired volume. Preferably the overfilling needs to be sufficient in size to compensate for the volume of spaces left between sintered particles of the filler material 16.
[0036] In step 108 the last layer 17 of the filler material 16 undergoes selective laser melting. Fig. 5 shows a fully dense skin 18 formed on the filler material 16 via the usage of selective laser melting. Densification is accomplished by melting and solidification of the last layer 17. A“fully dense skin” has from zero to less than five percent porosity. The selective laser melting of the last layer 17 of the filler material 16 forms a skin 18. The skin 18 is preferably porositiy-free. Such absence of porosity is important for the next step of consolidation by hot isostatic pressing (HIP). That is, (as is known in metal casting industry practice) surface pores are not generally closed by the HIP process, a solid skin ensures pores at and near to (e.g. within a layer thickness of) the surface get closed.
[0037] The connection point 19 between the skin 18 and the remainder of the gas turbine engine component should not adversely affect the gas turbine engine component 10. The skin 18 and the adjoining melted area of the gas turbine engine component 10 should be no thicker than 0.5 mm, preferably no thicker than 0.1 mm and most advantageously within the range of 25 to 100 microns. Preferably the connection 19 is of sufficient strength to survive the process that it undergoes in step 110. However, the thickness of the skin 18 is of no consequence so long as the process discussed below causes the skin 18 to collapse on the remainder of the material of the gas turbine engine component 10.
[0038] In step 110, the filler material 16 of the gas turbine engine component 10 goes through hot isotactic pressing. This is to fully compress the filler material 16 within the cavity 14. Fig. 6 shows the gas turbine engine component 10 after the application of hot isotactic pressing. After the hot isotactic pressing that occurs in step 110, the repaired gas turbine engine component can undergo machining and inspection prior to being returned to the gas turbine engine.
[0039] Now turning to Fig. 7, a flow chart setting forth an alternative method for repairing a gas turbine engine component is shown. The flow chart and method set forth in Fig. 7 is similar to the method set forth above. The differences address a potential issue of the embodiment found in Figure 5. This is accomplished in that the connection 19 could be too fragile if the filler material 16 in this situation was that same as that of the remainder of the gas turbine engine component 10. So in this case the filler material 16 is chosen to be matching or a superalloy material, but the final layer 17 (Figure 11) is chosen to be a ductile alloy. This can occur in situations where the gas turbine engine component 10 is made of a cast nickel based superalloy such as Inconel 939, Inconel 738 Rene 80, CM 247 or case cobalt based alloy such as earlier described.
[0040] Fig. 8 shows the gas turbine engine component 10 having a defective portion 12. The defective portion 12 may a pre-existing flaw that exists in the gas turbine engine component 10, or may be the result of damage that occurred during the operation of the gas turbine engine. Upon discovery of the defective portion, in step 202 the defective portion 12 may be removed from the gas turbine engine component 10. This may be achieved via excavation of the portion using grinding, ball mill machining, electro discharge machining, air arc gouging and other similar commercial methods of material removal.
[0041] Fig. 9 shows the gas turbine engine component 10 with the defective portion 12 removed. Shown is the cavity 14 that previously held the defective portion 12 of the gas turbine engine component 10.
[0042] In step 204, filler material 16 is placed in the cavity 14. Fig. 10 shows a gas turbine engine component 10 having the filler material 16 placed in the cavity 14 where the defective portion 12 was removed. The filler material 16 may be a matching or alternate superalloy in this embodiment. In this alternate embodiment the last layer is deposited of a sacrificial ductile alloy that is readily welded to the substrate. For example the ductile alloy may be Inconel alloy 600, 625, 617, 82 or Haynes alloy X, 282 , a stainless steel such as 304, 316, 347 or a ductile cobalt based alloy such as Haynes 188, L-605, or Mar-M 918. In step 206, the filler material 16 undergoes selective laser sintering. The selective laser sintering fills and preferably overfills the cavity 14 with filler material 16.
[0043] Steps 204 and 206 preferably involve placing and sintering material in relatively thin, e.g. 20 to 100 micron thick, layers and repeating such layered placement and sintering to build up and fill the cavity 14.
[0044] In step 208 the last layer 17 of the final ductile layer of filler material 16 undergoes selective laser melting and full densification. Fig. 11 shows a skin 18 formed on the filler material 16 via the usage of selective laser melting of the ductile alloy. The selective laser melting of the last layer 17 of the filler material 16 forms a skin 18. The skin 18 is preferably porositiy-free.
[0045] The connection point 19 between the skin 18 and the remainder of the gas turbine engine component should not adversely affect the gas turbine engine component 10. The skin 18 and the adjoining melted area of the gas turbine engine 10 should be no thicker than 0.5 mm, preferably no thicker than 0.1 mm and most advantageously within the range of 25 to 100 microns. Preferably the connection 19 is of sufficient strength to survive the process that it undergoes in step 210. In step 210, the filler material 16 of the gas turbine engine component 10 goes through hot isotactic pressing. This is to fully compress the filler material 16 within the cavity 14. Fig. 12 shows the gas turbine engine component 10 after the application of hot isotactic pressing.
[0046] After the hot isotactic pressing that occurs in step 210, in step 212 the skin 18 that is formed from the ductile alloy may be removed. The repaired gas turbine engine component 10 can undergo machining and inspection prior to being returned to the gas turbine engine.
[0047] While embodiments of the present disclosure have been disclosed in exemplary forms, it will be apparent to those skilled in the art that many modifications, additions, and deletions can be made therein without departing from the spirit and scope of the invention and its equivalents, as set forth in the following claims.

Claims

What is claimed is:
1. A method for repairing a gas turbine engine component (10) comprising:
removing a defective portion (12) of the gas turbine engine component (10), wherein the defective portion (12) is made of a gamma prime strengthened superalloy material;
depositing filler material (16) in a cavity (14) left by the removal of the defective portion;
sintering the filler material (16) with selective laser sintering;
depositing the filler material (16) and selective layer sintering the filler material (16) until the cavity (14) is filled or overfilled with the filler material (16); sealing a last layer (17) of the filler material (16) by forming a skin (18) over the filler material (16) with selective laser melting; and
hot isotactic pressing the filler material (16).
2. The method of claim 1, further comprising confirming that the defective portion (12) was removed using non-destructive inspection.
3. The method of claim 1, wherein the gamma prime strengthened superalloy material is selected from a group consisting of cobalt based superalloy materials, nickel based superalloy materials, nickel-cobalt based superalloy materials, iron based superalloy materials, iron-nickel based superalloy materials and oxide dispersion strengthened superalloys.
4. The method of claim 1, wherein the skin formed (18) over the filler material (16) is porosity free.
5. The method of claim 4, wherein a thickness of the skin (18) is less than 0.5 mm.
6. The method of claim 5, wherein the thickness is in a range of 25 to 100 microns.
7. The method of claim 1, wherein the filler material (16) is gamma prime
strengthened superalloy material.
8. A method for repairing a gas turbine engine component (10) comprising:
removing a defective portion (12) of the gas turbine engine component (10), wherein the defective portion (12) is made of a gamma prime strengthened superalloy material;
depositing filler material (16) in a cavity (14) left by the removal of the defective portion (12);
sintering the filler material (16) with selective laser sintering;
depositing the filler material (16) and selective layer sintering the filler material until the cavity (14) is filled or overfilled with the filler material (16);
sealing a last layer (17) of the filler material (16) by melting and fully densifying a ductile alloy skin (18) over the last layer (17) of the filler material (16); and
hot isotactic pressing the filler material (16).
9. The method of claim 8, wherein the ductile alloy skin (18) is removed after the step of hot isotactic pressing.
10. The method of claim 8, further comprising confirming that the defective portion (12) was removed using non-destructive inspection.
11. The method of claim 8, wherein the gamma prime strengthened superalloy material is selected from a group consisting of cobalt based superalloy materials, nickel based superalloy materials, nickel-cobalt based superalloy materials, iron based superalloy materials, iron-nickel based superalloy materials and oxide dispersion strengthened superalloys.
12. The method of claim 8, wherein the filler material (16) is gamma prime strengthened superalloy material.
13. The method of claim 8, wherein the sacrificial ductile material is selected from a group consisting of wrought nickel based or cobalt based alloys.
14. A method for repairing a gas turbine engine component (10) comprising:
removing a defective portion (12) of the gas turbine engine component (10), wherein the defective portion (12) is made of a gamma prime strengthened cobalt based superalloy material;
depositing filler material (16) in a cavity (14) left by the removal of the defective portion (12);
sintering the filler material (16) with selective laser sintering;
depositing the filler material (16) and selective layer sintering the filler material (16) until the cavity is filled or overfilled with the filler material (16);
sealing a last layer (17) of the filler material (16) by forming a skin over the filler material (16) with selective laser melting; and
hot isotactic pressing the filler material (16).
15. The method of claim 14, further comprising confirming that the defective portion (12) was removed using non-destructive inspection.
16. The method of claim 14, wherein the skin (18) formed over the filler material (16) is porosity free.
17. The method of claim 14, wherein a thickness of the skin (18) is less than 0.5 mm.
18. The method of claim 17, wherein the thickness is in a range of 25 to 100 microns.
19. The method of claim 14, wherein the filler material (16) is gamma prime strengthened superalloy material.
PCT/US2017/061722 2017-11-15 2017-11-15 Method of repairing gamma prime strengthened superalloys WO2019099000A1 (en)

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