EP1666618B2 - Superalliage à base Ni et son utilisation comme disques, arbres et rotors de turbines à gaz - Google Patents

Superalliage à base Ni et son utilisation comme disques, arbres et rotors de turbines à gaz Download PDF

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EP1666618B2
EP1666618B2 EP06001713.4A EP06001713A EP1666618B2 EP 1666618 B2 EP1666618 B2 EP 1666618B2 EP 06001713 A EP06001713 A EP 06001713A EP 1666618 B2 EP1666618 B2 EP 1666618B2
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percent
article
mass
temperature
solution
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English (en)
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EP1666618B1 (fr
EP1666618A1 (fr
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David Paul Mourer
Daniel Gustov Backman
Paul Leray Reynolds
Timothy Paul Gabb
Eric Scott Huron
Kenneth Rees Bain
John Joseph Schirra
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General Electric Co
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General Electric Co
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    • C—CHEMISTRY; METALLURGY
    • C22—METALLURGY; FERROUS OR NON-FERROUS ALLOYS; TREATMENT OF ALLOYS OR NON-FERROUS METALS
    • C22C—ALLOYS
    • C22C19/00—Alloys based on nickel or cobalt
    • C22C19/03—Alloys based on nickel or cobalt based on nickel
    • C22C19/05—Alloys based on nickel or cobalt based on nickel with chromium
    • C22C19/051—Alloys based on nickel or cobalt based on nickel with chromium and Mo or W
    • C22C19/056—Alloys based on nickel or cobalt based on nickel with chromium and Mo or W with the maximum Cr content being at least 10% but less than 20%
    • 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
    • B22F3/24—After-treatment of workpieces or articles
    • 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
    • C22F—CHANGING THE PHYSICAL STRUCTURE OF NON-FERROUS METALS AND NON-FERROUS ALLOYS
    • C22F1/00—Changing the physical structure of non-ferrous metals or alloys by heat treatment or by hot or cold working
    • C22F1/10—Changing the physical structure of non-ferrous metals or alloys by heat treatment or by hot or cold working of nickel or cobalt or alloys based thereon
    • 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
    • B22F3/24—After-treatment of workpieces or articles
    • B22F2003/248—Thermal after-treatment

Definitions

  • This invention relates to a superalloy having nickel as the major component, and, more particularly, to such a superalloy particularly useful in the production of gas turbine disks, impellers, and shafts by powder metallurgy techniques.
  • a gas turbine (jet) engine air is drawn into the front end of the engine, compressed by a shaft-mounted compressor disk, and mixed with fuel. The mixture is ignited, producing a hot exhaust gas that is passed through a turbine which provides the power to the compressor, and then exhausted rearwardly to drive the engine and the aircraft, in which it is mounted, forwardly.
  • the turbine In the axial flow jet engine, the turbine has a turbine disk which is mounted to a drive shaft, and turbine blades extending from the periphery of the turbine disk.
  • the compressor disk is mounted to its shaft, which is driven by the turbine shaft.
  • Turbine disk must carry high multiaxial loads in tension, and must exhibit good creep resistance and dwell fatigue capability as well as good fracture toughness.
  • Turbine disks for use at moderately high temperatures have in the past typically been forged, which tends to produce a degree of anisotropy in the disk. As the operating temperatures have been increased through improvements in alloy compositions, other fabrication techniques have been developed.
  • the alloy material of construction is provided in the form of fine powders. These powders are compacted together in the form of the turbine disk or shaft, usually by extrusion and isothermal forging, and then heat treated and final machined as necessary.
  • the final article is largely isotropic due to the use of the powders, and has properties determined by the composition of the powder particles and the heat treatment.
  • US-A-5120373 discloses a process for producing a fine grain forged superalloy article having a high yield strength at intermediate temperatures.
  • a preferred starting composition comprises, by weight, 15% Cr, 13.6%Co, 4.1% Mo, 4.6% Ti, 2.2% Al, 0.01 % C, 0.007%B, 0.07%Zr, balance Ni.
  • This material is forged at a temperature above the gamma prime solvus temperature and a truestrain of at least 0.5. Alternatively the material may be forged below the gamma prime solvus.
  • EP-A-0373298 discloses a composition consisting of, in weight percent, 12-18 Co, 7-13 Cr, 2-4 Mo, 0-1.0 W, 2.0-5.0 Al, 5.0-7.0 Ti, 2.0-3.2 Ta, 1.0-1.7 Nb, 0-0.75 Hf, 0-0.1 Zr, 0.0-2.0 V, 0-0.2 C, 0-0.1 B, 0-1.0 Re, 0-0.1 Y, balance Ni.
  • EP-A-0849370 discloses a nickel base superalloy article having a machined surface having a composition in weight percent, 1.2-3.5 Al, 3.0-7.0 Ti, 12.0-20 Cr, 2.0-3.9 Mo, 10.0-20.0 Co, 1-4.5 W, 0.005-0.025 C, 0.005-0.05 B, 0.01-0.1 Zr, 0-0.005 Mg, 0-1.0 Ta, 0-1.0 Nb, 0-2.0 Fe, 0-0.3 Hf, 0-0.02 Y, 0-0.1 V, 0-1.0 Re, balance Ni.
  • the article has a plurality of discrete carbides essentially free from molybdenum, for increased fatigues strength.
  • the present invention provides compositions of matter, articles using the compositions of matter, and processing methods for the compositions of matter that achieve improved combinations of properties in conditions experienced in aircraft gas turbine disk and shaft applications. Both dwell fatigue crack growth rate and time to creep specific amounts or elongation are improved as compared with other alloys used for these applications. This combination of improved properties is particularly advantageous for use in aircraft engines which are not operated at the temperatures required for advanced military fighter engines but which spend long periods at moderately elevated temperature in cruise conditions.
  • the selected compositions reflect careful balancing of the amounts of both the major and minor elements.
  • a composition of matter consists of, in weight percent, from 16 percent to 20 percent cobalt, from 11 percent to 15 percent chromium, from 2 percent to 4 percent tantalum, from 0.5 to 3 percent tungsten, from 3 to 5 percent molybdenum, from 0.015 to 0.15 percent zirconium, from 1 to 3 percent niobium, from 2.6 to 4.6 percent titanium, from 2.6 to 4.6 percent aluminum, from 0 to 2.5 percent rhenium, from 0 to 2 percent vanadium, from 0 to 2 percent iron, from 0 to 2 percent hafnium, from 0 to 0.1 percent magnesium, from 0.015 to 0.1 percent carbon, from 0.015 percent to 0.045 percent boron, balance nickel and impurities.
  • the ratio (percent zirconium + percent boron)/percent carbon is greater than 1.0.
  • compositions of the invention are prepared in powder form, and processed into articles by combinations of extrusion, hot isostatic pressing, isothermal forging, heat treating, and other operable techniques.
  • the preferred articles made with these compositions are turbine and compressor disks and shafts, and compressor impellers for gas turbine engines.
  • the articles may be heat treated, either by solution treating and ageing or by solution treating followed by a controlled cooling to below the solvus temperature to control residual stresses.
  • the articles made according to the invention exhibit a combination of low dwell fatigue crack growth rate and long creep times that are unexpectedly improved over prior materials used for the same applications.
  • compositions, articles, and methods of the present invention result in improved dwell fatigue crack growth rate and creep properties, while retaining acceptable density and other physical and mechanical properties.
  • This combination of properties is particularly advantageous for use in turbine disk applications in advanced civilian aircraft engines, where the engine has an extended operating cycle at elevated temperature, but where the temperature requirements of the engine are not as great as in military aircraft.
  • Figure 1A shows a turbine disk 20
  • Figure 1B shows a compressor impeller 22
  • Figure 2 shows a turbine shaft 24 used in a gas turbine engine, each of which may be made by the approach of the invention.
  • a compressor disk has an appearance which is generally similar to that of a turbine disk
  • a compressor shaft has an appearance which is generally similar to that of a turbine shaft.
  • Figure 3 depicts a method of fabricating articles such as those of Figures 1 and 2 .
  • a metallic composition of matter is furnished, numeral 30.
  • the composition of matter of the present invention is, in weight percent, from 16 percent to 20 percent cobalt, from 11 percent to 15 percent chromium, from 2 percent to 4 percent tantalum, from 0.5 to 3 percent tungsten, from 3 to 5 percent molybdenum, from 0.015 to 0.15 percent zirconium, from 1 to 3 percent niobium, from 2.6 to 4.6 percent titanium, from 2.6 to 4.6 percent aluminum, from 0 to 2.5 percent rhenium, from 0 to 2 percent vanadium, from 0 to 2 percent iron, from 0 to 2 percent hafnium, from 0 to 0.1 percent magnesium from 0.015 to 0.1 percent carbon, from 0.015 percent to 0.045 percent boron, balance nickel and impurities.
  • This alloy composition produces a gamma/gamma prime microstructure, which may be controlled through heat treatments, with minor amounts of other phases present such as borides and carbides.
  • the gamma prime phase is present in an amount, based on calculation, of from 47 to 55 volume percent of the total volume of the material, in order to produce the desirable properties of the alloy.
  • the types and amounts of the elements in the alloy composition are chosen in cooperation with each other to achieve the desired properties, based upon testing and the analysis undertaken by the inventors. Due to the interaction between the elements, the experimental compositions defined the trends for alloying, but only limited ranges of alloy compositions exhibit the final effects of compositional influences, microstructures, and resulting properties. Together the alloying trends and the absolute elemental levels define the preferred ranges of compositions. The effects of individual elements and the results of their amounts in the alloys falling outside the indicated ranges may be summarized as follows.
  • the cobalt level is selected to control the gamma prime solvus temperature. Increasing amounts of cobalt lower the gamma prime solvus temperature, which is desirable to achieve a large processing temperature range and reduce the stresses induced by controlled cooling or quenching of the alloy used to define a portion of the gamma prime distribution and the preferred combination of mechanical properties. If the amount of cobalt is substantially less than that indicated, the gamma prime solvus temperature is too high and there is a risk of incipient melting or thermally induced porosity. If the cobalt content is substantially greater than that indicated, the alloy has an undesirably higher elemental cost.
  • chromium is beneficial to oxidation resistance, corrosion resistance, and fatigue crack growth resistance. If the amount of chromium present is substantially less than that indicated, these properties may suffer. If it is substantially more than that indicated, there may be alloy, chemical, or phase instability during extended exposure to elevated temperatures, and creep performance suffers.
  • the control of the refractory elements tantalum, tungsten, niobium, and molybdenum is important to achieving the balance required in the alloy and articles of the invention.
  • Tantalum whose presence and percentage content of tantalum is important to achieving the beneficial results obtained for the alloys of the invention, primarily enters the gamma-prime phase and has the effect of improving the stability of the gamma-prime phase and improving the creep resistance and fatigue crack growth resistance of the alloy. If the tantalum content is substantially lower than these amounts, the creep life of the alloy is reduced and the dwell fatigue crack growth resistance is insufficient. Increasing the tantalum substantially above the indicated amounts has the undesirable effect of raising the gamma-prime solvus temperature so as to reduce the processibility of the alloy and increase its density.
  • Tungsten and niobium are two relatively dense elements which function together to achieve synergistic positive results with respect to creep capability.
  • Figure 4 shows the time for a standard tensile specimen to creep to 0.2 percent elongation at 649°C (1200°F) and under a load of 793 MPa (115,000 pounds per square inch), With less than 0.5 weight percent tungsten and less than 1 weight percent niobium, or with one or the other of the two elements present but not both, the creep properties are relatively poor. If both tungsten and niobium are present above these indicated minimum limits, the creep properties are markedly better.
  • Tungsten enters the matrix as a solid-solution strengthening element, and also aids in forming gamma prime precipitates. If the amount of tungsten is substantially less than that indicated, the creep properties may be insufficient. However, tungsten is relatively dense and also can lead to notch sensitivity and chemical instability. If the amount of tungsten is substantially greater than that indicated, the density of the alloy is too high, and, in addition, notch sensitivity is enhanced and chemical instability is of concern.
  • niobium is relatively dense and also can lead to notch sensitivity, chemical instability, and loss of dwell fatigue crack growth capability. If the amount of niobium is substantially greater than that indicated, the density of the alloy is too high, and, in addition, notch sensitivity is enhanced and chemical instability and reduced dwell fatigue crack growth capability are of concern.
  • Molybdenum is another relatively dense refractory element that partitions primarily to the gamma phase and has a beneficial effect on creep capability. If the amount of molybdenum is substantially less than 3 weight percent, the creep capability of the material may be reduced below desirable levels. If the amount of molybdenum is greater than about 5 weight percent, alloy stability is reduced and alloy density is increased above the desired level.
  • Titanium is a relatively light element and therefore may be added more freely to the alloy, from a density standpoint, to contribute to gamma prime formation. If titanium is present in an amount substantially less than that indicated, the tensile and dwell fatigue crack growth properties may be insufficient. If titanium is present in an amount substantially greater than that indicated, the heat treat window may be unacceptably reduced because the gamma prime solvus temperature is raised excessively. Substantially greater titanium levels may also stabilize or produce undesirable phases such as eta phase, which ties up the titanium and prevents it from participating in the production of the desired gamma prime microstructure.
  • Aluminum is present to contribute to gamma prime phase formation and to promote gamma prime phase stability.
  • Aluminum is the lowest-density gamma prime forming element and offsets the presence of higher-density elements. If aluminum is present in an amount substantially less than or greater than that indicated, then too little or too much of the gamma prime phase is present, and the stability of the alloy is adversely affected.
  • Carbon is present to aid in controlling grain size of the alloy. If the carbon content is substantially less than that indicated, the grain size of the alloy tends to grow too large, particularly during supersolvus processing. However, if the carbon content is substantially greater than that indicated, the carbon may have an adverse effect on the fracture properties of the alloy through premature failure. The higher carbon content also adversely affects the dwell fatigue crack growth resistance and creep capability.
  • Boron in moderate amounts improves the dwell fatigue crack growth resistance. If the boron is substantially less than that indicated, the alloy has insufficient dwell fatigue crack growth resistance. However, boron in an amount substantially greater than that indicated tends to cause residual porosity or thermally induced porosity and incipient melting during processing, and to reduce creep capability.
  • Zirconium is present in an amount of from 0.015 percent to 0.15 percent, more preferably from 0.35 to 0.055 percent, and most preferably from 0.04 to 0.05 percent.
  • the presence of zirconium in controlled small amounts improves the elongation and ductility of the alloy, and also reduces the crack growth rate.
  • Zirconium in amounts substantially in excess of the indicated levels tends to increase the creep rate of the alloy.
  • rhenium in an amount up to 2.5 percent by weight, magnesium in an amount up to 0.1 percent by weight, vanadium in an amount up to 2 percent by weight, iron in an amount up to 2 percent by weight, and hafnium in an amount up to 2 percent by weight may be present without adversely affecting the properties.
  • the hafnium may improve the dwell fatigue crack growth rate but with a slight negative effect on low cycle fatigue.
  • alloy ME1-16 has a composition of, in weight percent, 18.2 percent cobalt, 13.1 percent chromium, 2.7 percent tantalum, 1.9 percent tungsten, 3.8 percent molybdenum, 0.050 percent zirconium, 1.4 percent niobium, 3.5 percent titanium, 3.5 percent aluminum, 0.030 percent carbon, 0.030 percent boron, balance nickel and impurities.
  • Another specific most preferred alloy has a composition, in weight percent, of 20 percent cobalt, 13 percent chromium, 2 percent tantalum, 2 percent tungsten, 3.8 percent molybdenum, 0.050 percent zirconium, 1.2 percent niobium, 3.7 percent titanium, 3.7 percent aluminum, 0.05 percent carbon, 0.03 percent boron, balance nickel and impurities.
  • compositions are a result of the selection of the combination of elements, not any one element in isolation.
  • the more preferred and most preferred compositions yield progressively improved results than the broad composition within the operable range, but it is also possible to attain improved results by combining the narrowed composition ranges of some elements producing improved results with the broader composition ranges of other elements.
  • the alloy composition is formed into a powder, numeral 32, by any operable technique. Gas or vacuum atomization is preferred.
  • the powder particles are preferably finer than -60 mesh, and most preferably -140 mesh or -270 mesh.
  • the powder is consolidated to a billet or forging preform shape and then subsequently deformed to a final shape, numeral 34.
  • the preferred approach to consolidation is extrusion processing at an extrusion temperature of from 1010°C (1850°F) to 1107°C (2025°F), and a 3:1 to 6:1 extrusion ratio.
  • the alloy is deformed to a shaped contour oversize to, but approximating the outline of, the final part.
  • the deformation step is preferably accomplished by isothermal forging in a strain-controlled mode.
  • the consolidation, deformation, and a subsequent supersolvus solution heat treatment are preferably selected to yield a grain size of from ASTM 2 to ASTM 8, preferably from ASTM 5 to ASTM 8.
  • the consolidation, deformation, and a subsequent subsolvus solution heat treatment are selected to yield a grain size of from ASTM 9 to ASTM 12, preferably from ASTM 10 to ASTM 12.
  • the extruded article is heat treated, numeral 36, to produce the desired microstructure.
  • the article is solution heat treated by heating to a supersolvus temperature, such as from 1149°C (2100°F) to 1218°C (2225°F) for a period of time sufficient that the entire article reaches this temperature range.
  • the solution-treated article is quenched (cooled) to room temperature by a fan air cool, optionally followed by an oil quench.
  • the solution-treated-and-quenched article is then aged by reheating to a temperature below the solvus temperature, preferably from 732°C (1350°F) to 816°C (1500°F) for a time of about 8 hours.
  • the article may be stress relieved by heating it to a stress-relieving temperature of from 816°C (1500°F) to 982°C (1800°F), most preferably 843°C (1550°F) for 4 hours, either after the quenching step and before the aging step, or after the final age step.
  • a stress-relieving temperature of from 816°C (1500°F) to 982°C (1800°F), most preferably 843°C (1550°F) for 4 hours, either after the quenching step and before the aging step, or after the final age step.
  • the article is solution treated at a partial subsolvus solution-treating temperature of from 1093°C (2000°F) to 1149°C (2100°F), quenched as described above and aged, or cooled, stress relieved and aged, as described above.
  • the article is slow cooled from a supersolvus solution temperature at rates of less than 260°C (500°F) per hour to a subsolvus temperature.
  • the article is then quenched as described above and aged, or stress relieved and aged, as described above.
  • Figure 6 also illustrates data for the time for reach 0.2 percent creep when measured at 649°C (1200°F) and a stress of 793 MPa (115,000 pounds per square inch).
  • compositions of the present invention achieve significantly improved dwell fatigue crack growth rates and improved creep times, as compared with conventional alloys.
  • data is presented for IN100 and Rene 88DT, standard disk and shaft alloys.
  • Alloy ME 1-16 and ME2 are within the scope of the present invention, alloy ME1-12 is outside the scope of the invention.
  • Alloy CH98 is the preferred composition disclosed in US Patent 5,662,749 .
  • the alloys of the present invention achieve an improvement of approximately a factor of 50 over IN100 in creep life and approximately a factor of 200 over Rene 88DT in dwell fatigue crack growth rate.
  • the alloys of the present invention have about the same dwell fatigue crack growth performance as alloy CH98, and exhibit substantially improved creep life over alloy CH98.
  • the present alloys provide a level of enhanced performance for both dwell fatigue crack growth rate and time to creep that is desirable for articles such as gas turbine disks and shafts that are subjected to both types of loading during service.

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Claims (20)

  1. Composition de matière constituée, en pour-cent en poids, de 16 à 20 pour-cent de cobalt, de 11 à 15 pour-cent de chrome, de 2 à 4 pour-cent de tantale, de 0,5 à 3 pour-cent de tungstène, de 3 à 5 pour-cent de molybdène, de 0,015 à 0,15 pour-cent de zirconium, de 1 à 3 pour-cent de niobium, de 2,6 à 4,6 pour-cent de titane, de 2,6 à 4,6 pour-cent d'aluminium, de 0 à 2,5 pour-cent de rhénium, de 0 à 2 pour-cent de vanadium, de 0 à 2 pour-cent de fer, de 0 à 2 pour-cent d'hafnium, de 0 à 0, 1 pour-cent de magnésium, de 0,015 à 0,1 pour- cent de carbone, de 0,015 à 0,045 pour-cent de bore, le complément étant du nickel et des impuretés ; dans laquelle le rapport (pour-cent de zirconium + pour-cent de bore)/ pour-cent de carbone est supérieur à 1.
  2. Composition de matière selon la revendication 1, dans laquelle la somme de tungstène plus niobium est comprise entre 2,2 pour-cent et 4 pour-cent.
  3. Composition de matière de la revendication 1, dans laquelle la composition est constituée, en pour-cent en poids, de 18,2 pour-cent environ de cobalt, 13,1 pour-cent environ de chrome, 2,7 pour-cent environ de tantale, 1,9 pour-cent environ de tungstène, 3,8 pour-cent environ de molybdène, 0,050 pour-cent environ de zirconium, 1,4 pour-cent environ de niobium, 3,5 pour-cent environ de titane, 3,5 pour-cent environ d'aluminium, 0,030 pour-cent environ de carbone, 0,030 pour-cent environ de bore, le complément étant du nickel et des impuretés.
  4. Composition de matière selon la revendication 1, dans laquelle la composition est constituée, en pour-cent en poids, de 20 pour-cent environ de cobalt, 13 pour-cent environ de chrome, 2 pour-cent environ de tantale, 2 pour-cent environ de tungstène, 3,8 pour-cent environ de molybdène, 0,050 pour- cent environ de zirconium, 1,2 pour-cent environ de niobium, 3,7 pour-cent environ de titane, 3,7 pour-cent environ d'aluminium, 0,05 pour-cent environ de carbone, 0,03 pour-cent environ de bore, le complément étant du nickel et des impuretés.
  5. Article présentant une composition selon la revendication 1.
  6. Article selon la revendication 5, dans lequel l'article comprend une masse de poudres compactées.
  7. Article selon la revendication 6, dans lequel l'article présente une taille de grain de ASTM 2 à ASTM 8.
  8. Article selon la revendication 5, dans lequel l'article présente une taille de grain de ASTM 9 à ASTM 12.
  9. Article selon la revendication 5, dans lequel l'article est sélectionné à partir du groupe constitué par un disque de turbine (20), un arbre de turbine (24), un disque de compresseur (20), un arbre de compresseur (24), et un rotor de compresseur (22).
  10. Procédé de préparation d'un article, comprenant les étapes de Fourniture d'une masse de poudres compactées présentant une composition constituée, en pour-cent en poids, de 16 à 20 pour-cent de cobalt, de 11 à 15 pour-cent de chrome, de 2 à 4 pour-cent de tantale, de 0,5 à 3 pour- cent de tungstène, de 3 à 5 pour-cent de molybdène, de 0,015 à 0,15 pour-cent de zirconium, de 1 à 3 pour-cent de niobium, de 2,6 à 4,6 pour-cent de titane, de 2,6 à 4,6 pour-cent d'aluminium, de 0 à 2,5 pour-cent de rhénium, de 0 à 2 pour-cent de vanadium, de 0 à 2 pour-cent de fer, de 0 à 2 pour-cent de hafnium, de 0 à 0,1 pour-cent de magnésium, de 0,015 à 0,1 pour-cent de carbone, de 0,015 à 0,045 pour-cent de bore, le complément étant du nickel et des impuretés ; dans lequel le rapport (pour-cent de zirconium + pour-cent de bore)/ pour-cent de carbone est supérieur à 1.
    traitement thermique de la masse par les étapes de
    traitement en solution de la masse à une température de traitement supérieure à sa température de limite de solubilité,
    et refroidissement de la masse traitée en solution à une température inférieure à sa température de limite de solubilité.
  11. Procédé selon la revendication 10, dans lequel l'étape de traitement thermique comprend une étape supplémentaire, après l'étape de refroidissement, de maturation de la masse traitée en solution et recuite à une température de maturation inférieure à sa température de limite de solubilité.
  12. Procédé selon la revendication 11, dans lequel l'étape de maturation comporte l'étape de chauffage de la masse à une température de maturation de 732°C (1350°F) à 816°C (1500°F).
  13. Procédé selon la revendication 11, comportant une étape supplémentaire, après l'étape de refroidissement, de libération de contraintes de l'article par chauffage de l'article à une température de libération de contraintes de 816°C (1500°F) à 982°C (1800°F).
  14. Procédé selon la revendication 10, dans lequel l'étape de traitement en solution comporte l'étape de chauffage de la masse à une température de traitement en solution de 1149°C (2100°F) à 1218°C (2225°F).
  15. Procédé selon la revendication 10, dans lequel l'étape de traitement en solution comporte l'étape de chauffage de la masse à une température de traitement en solution inférieure à la limite de solubilité partielle de 1093°C (2000°F) à 1149°C environ (2100°F).
  16. Procédé selon la revendication 15, dans lequel l'étape de traitement thermique comporte une étape supplémentaire, après l'étape de refroidissement,
    de maturation de la masse traitée en solution et refroidie au-dessous de la limite de solubilité partielle à une température de maturation inférieure à sa limite de solubilité.
  17. Procédé selon la revendication 16, dans lequel l'étape de maturation comprend l'étape de
    chauffage de la masse traitée en solution et refroidie au-dessous de la limite de solubilité partielle à une température de maturation de 732°C (1350°F) à 816°C (1500°F).
  18. Procédé selon la revendication 16, comportant une étape supplémentaire, après l'étape de refroidissement, de libération de contraintes de l'article à une température de libération de contraintes de 816°C (1500°F) à 982°C (1800°F).
  19. Composition de matière selon la revendication 1, dans laquelle la teneur en tungstène est de 1,4 à 3 pour-cent en poids.
  20. Article selon la revendication 6, dans lequel la teneur en tungstène est de 1,4 à 3 pour-cent en poids.
EP06001713.4A 2000-10-04 2000-10-04 Superalliage à base Ni et son utilisation comme disques, arbres et rotors de turbines à gaz Expired - Lifetime EP1666618B2 (fr)

Priority Applications (2)

Application Number Priority Date Filing Date Title
DE60041936T DE60041936D1 (de) 2000-10-04 2000-10-04 Ni-basis-Superlegierung und ihre Verwendung als Gasturbinen-Scheiben, -Wellen und -Laufräder
EP06001713.4A EP1666618B2 (fr) 2000-10-04 2000-10-04 Superalliage à base Ni et son utilisation comme disques, arbres et rotors de turbines à gaz

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
EP06001713.4A EP1666618B2 (fr) 2000-10-04 2000-10-04 Superalliage à base Ni et son utilisation comme disques, arbres et rotors de turbines à gaz
EP00308759A EP1195446A1 (fr) 2000-10-04 2000-10-04 Superalliage à base Ni et son utilisation comme disques, arbres et rotors de turbines à gaz

Related Parent Applications (1)

Application Number Title Priority Date Filing Date
EP00308759A Division EP1195446A1 (fr) 2000-10-04 2000-10-04 Superalliage à base Ni et son utilisation comme disques, arbres et rotors de turbines à gaz

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EP1666618A1 EP1666618A1 (fr) 2006-06-07
EP1666618B1 EP1666618B1 (fr) 2009-04-01
EP1666618B2 true EP1666618B2 (fr) 2015-06-03

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EP06001713.4A Expired - Lifetime EP1666618B2 (fr) 2000-10-04 2000-10-04 Superalliage à base Ni et son utilisation comme disques, arbres et rotors de turbines à gaz
EP00308759A Ceased EP1195446A1 (fr) 2000-10-04 2000-10-04 Superalliage à base Ni et son utilisation comme disques, arbres et rotors de turbines à gaz

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EP1666618B1 (fr) 2009-04-01
EP1195446A1 (fr) 2002-04-10
EP1666618A1 (fr) 2006-06-07

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