EP2944704B1 - Composition d'alliage de nickel - Google Patents

Composition d'alliage de nickel Download PDF

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Publication number
EP2944704B1
EP2944704B1 EP15163869.9A EP15163869A EP2944704B1 EP 2944704 B1 EP2944704 B1 EP 2944704B1 EP 15163869 A EP15163869 A EP 15163869A EP 2944704 B1 EP2944704 B1 EP 2944704B1
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EP
European Patent Office
Prior art keywords
alloy
gas turbine
titanium
chromium
aluminium
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Not-in-force
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EP15163869.9A
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German (de)
English (en)
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EP2944704A1 (fr
Inventor
Bryce Conduit
Gareth Conduit
Howard Stone
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Rolls Royce PLC
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Rolls Royce PLC
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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%
    • 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
    • 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
    • C22C19/055—Alloys based on nickel or cobalt based on nickel with chromium and Mo or W with the maximum Cr content being at least 20% but less than 30%
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/002—Wall structures

Definitions

  • the present invention relates to a nickel based alloy composition and a gas turbine engine component comprising a nickel based alloy.
  • Fig. 1 shows a high by pass ratio gas turbine engine 10.
  • the engine 10 comprises an air intake 11 and a propulsive fan 12 that generates two airflows A and B.
  • the gas turbine engine 10 comprises, in axial flow A, an intermediate pressure compressor 14, a high pressure compressor 16, a combustor 18, a high pressure turbine 20, an intermediate pressure turbine 22, a low pressure turbine 24 and an exhaust nozzle 26.
  • a nacelle 30 surrounds the gas turbine engine 10 and defines, in axial flow B, a bypass duct 13.
  • the combustor 20 is shown in further detail in Fig. 2 .
  • the combustor comprises a combustor casing 32, within which is located a metal combustor liner 34.
  • the combustor liner 34 is in turn covered in combustor liner tiles (not shown), which are made of a ceramic material. In use, air and fuel flow into the combustor 20, where the fuel is burned, producing hot combustion gases.
  • the combustor liner 34 must operate at high temperature in excess of 800°C (and perhaps as high as 900°C) for long periods of time. Higher combustion chamber temperature will result in higher thermal efficiencies of the gas turbine engine, and so this temperature must be made as high as possible through the use of high temperature alloys. High temperature alloys are also used in other parts of the engine, such as in the turbines and exhaust duct.
  • suitable alloys In order to permit operation at high temperatures, and provide a long service life, suitable alloys must also have a number of other properties, in addition to a high temperature capability. For example, they must have a high ultimate tensile strength, yield strength, stress rupture resistance, ductility, stability at high temperatures, resistance to thermal stresses, density and environmental resistance (e.g. resistance to hot corrosion and oxidation).
  • the "stability" of an alloy is normally understood to refer to the alloy's propensity to precipitate detrimental phases (i.e. an alloy having a high stability will have a low propensity to precipitate detrimental phases).
  • ⁇ phase An example of a detrimental phase is the sigma ( ⁇ ) phase, which can occur when the alloy is subjected to high temperatures for extended time periods (known as "dwell").
  • Table 1 below defines prior nickel based compositions suitable for use in combustor liners for gas turbine engines. All amounts are given in weight percentages.
  • the alloy composition prefferably has a low cost (in terms of the elemental cost of the alloy), and low density (particularly where the combustor is for use in aerospace gas turbine engines) and be suitable for low cost production methods such as casting, wrought processing, powder metallurgy or direct laser deposition. It is also desirable that the final alloy is highly weldable (i.e. does not micro-segregate when melted), particularly where the alloy is to be used in direct laser deposition. Alloys having the above properties, and therefore being suitable for use in gas turbine engine components such as combustor liners, are generally known in the art as “superalloys", and are sometimes also referred to as "high performance alloys".
  • the present invention seeks to provide an improved alloy composition and an improved gas turbine engine component which solves some or all of the above problems.
  • an alloy composition consisting essentially of, by weight per cent, between 18 and 20% chromium, between 8.5% and 12% cobalt, between 6.4 and 9.4% iron, up to 1.5% manganese, between 3.5 and 5.3% tungsten, up to 1.5% niobium, between 0.4% and 1 % tantalum, up to 1.5% silicon, between 1.5 and 2.3% aluminium, between 0.5 and 2.4% titanium, between 0.005 and 0.04% carbon, between 0.005 and 0.07% boron, and between 0.02 and 0.10% zirconium, optionally, up to 0.5% molybdenum, the balance being nickel and incidental impurities.
  • the alloy composition of the present invention has a high strength at temperatures at around 900°C, and is also relatively lightweight, resistant to environmental degradation, inexpensive and suitable for conventional forming and machining processes. These properties make the alloy particularly suitable for use in gas turbine engine components such as combustor linings.
  • a component of a gas turbine engine formed of an alloy in accordance with the first aspect of the invention.
  • a gas turbine engine comprising a component according to the second aspect of the invention.
  • the invention provides a gas turbine engine having one or more components capable of operation at high temperatures. Consequently, the gas turbine engine can operate at a higher thermal efficiency, and hence lower specific fuel consumption, or may have a longer service life between overhauls, resulting in a lower operating cost.
  • Fig. 1 shows a high by pass ratio gas turbine engine 10.
  • the engine 10 comprises an air intake 11 and a propulsive fan 12 that generates two airflows A and B.
  • the gas turbine engine 10 comprises, in axial flow A, an intermediate pressure compressor 14, a high pressure compressor 16, a combustor 18, a high pressure turbine 20, an intermediate pressure turbine 22, a low pressure turbine 24 and an exhaust nozzle 26.
  • a nacelle 30 surrounds the gas turbine engine 10 and defines, in axial flow B, a bypass duct 13.
  • the combustor 20 is shown in further detail in Fig. 2 .
  • the combustor comprises a combustor casing 32, within which is located a metal combustor liner 34.
  • the combustor liner 34 is in turn covered in combustor liner tiles (not shown), which are made of a ceramic material. In use, air and fuel flow into the combustor 20, where the fuel is burned, producing hot combustion gases.
  • Table 2 shows a compositional range of an alloy in accordance with the disclosure, which is suitable for one or more components of the gas turbine engine 10 (and particularly suitable for use as the material of the combustor liner 34): Table 2 wt.% target maximum minimum Ni Balance Balance Balance Cr 20.0-20.4 22 18.0 Co 9.3-11.3 12 8.5 Mo 0-0.2 0 0.5 Fe 7.4-8.4 9.4 6.4 Mn 0.2-0.4 1.5 0.0 W 4.2-4.6 5.3 3.5 Nb 1.0-1.2 1.5 0.0 Ta 0.6-0.8 1.0 0.4 Si 0.2-0.4 1.5 0 Al 1.8-2.0 2.3 1.5 Ti 1.6-1.8 2.4 0.5 C 0.015-0.025 0.04 0.005 B 0.015-0.025 0.04 0.005 Zr 0.05-0.07 0.10 0.02
  • Oxygen may also be present, in the form of surface oxides.
  • Other incidental impurities may also be present in the alloy. In general, other impurities should be kept to a minimum, in particular sulphurous based impurities.
  • compositions can be produced using the maximum and minimum elemental quantities described in table 2. It has been found that all compositions within the bounds of the maximum and minimum quantities of table 2 result in alloys which have acceptable properties for use as a combustor liner 34.
  • the narrower target compositional range shown in table 2 has improved qualities over the alloy compositions lying outside this range, but within the wider range shown in the maximum and minimum columns. These minimum and maximum amounts are based upon sensitivity studies using a computational materials prediction tool which show the deviation of predicted properties with changing composition.
  • Table 3 shows a nominal composition in accordance with the present disclosure (composition 1). The physical properties of this composition are described in further detail below.
  • Table 3 wt.% Composition Ni Bal. Cr 20.2 Co 10.3 Mo 0.1 Fe 7.9 Mn 0.3 W 4.4 Nb 1.1 Ta 0.7 Si 0.3 Al 1.9 Ti 1.7 C 0.015-0.025 B 0.015-0.025 Zr 0.06
  • oxygen may also be present, in the form of surface oxides.
  • impurities may also be present.
  • the composition comprises nickel, which forms a continuous matrix comprising a face centred cubic (FCC) nickel based austenitic phase containing solid solution elements.
  • the nickel based continuous matrix is known within the art as the "gamma ( ⁇ )" phase.
  • Some of the alloying elements form a primary strengthening phase known as a "gamma prime ( ⁇ ')” phase in an amount such that the volume fraction of the ⁇ ' is approximately 20% at 900°C.
  • the ⁇ ' phase has the general formula Ni 3 x, where x comprises elements selected from titanium, aluminium, tantalum and niobium, and usually comprises an ordered intermetallic L 12 crystal structure. Formation of the ⁇ ' phase occurs in the solid state as the supersaturated solid solution of ⁇ -nickel is cooled below its solvus temperature. Other elements (such as cobalt, iron, and tungsten) provide solid solution strengthening within the nickel matrix.
  • each of aluminium, titanium tantalum and niobium partition to the gamma phase.
  • the relative amounts of each of these elements may vary between compositions within the scope of the present disclosure, the total amount of elements that partition to the gamma prime phase is between 6.5 and 8.5 atomic per cent. This has been found to contribute to the desirable properties of the disclosed alloy, in particular, ultimate tensile strength.
  • the alloy includes the following refractory (i.e. high melting temperature) alloys, which offer significant strengthening in the alloy at the temperatures at which gas turbine engines operate: tungsten, molybdenum, niobium and tantalum. These refractory alloys are included in amounts between 1.8 and 3.8 atomic per cent.
  • the chromium present in the composition is required to maintain a protective oxide scale on the surface of the billet (or finished component), providing resistance to oxidation, type I and II type hot corrosion, and dwell fatigue crack resistance. If the chromium content is too high (i.e. significantly above (20%), then the formation of deleterious phases is encouraged, which will impair the mechanical properties of the alloy.
  • Cobalt is added to provide additional solid solution strengthening to the gamma matrix and reduce the stacking fault energy. Cobalt can be added in comparatively large quantities (up to 12%) due to its compatibility with nickel. However, too much cobalt (i.e. significantly more than 12%) will increase the propensity to form deleterious phases in the alloy at the temperatures at which gas turbine engine components typically operate.
  • Iron has good solubility within the gamma phase, and is added as a solid solution strengthener. It has the additional benefits that it is low cost and decreases the density of the alloy, resulting in a lightweight, low cost alloy. However, an iron content that is too high (i.e. significantly above 10%) will promote the formation of the undesirable Laves phase at the temperatures at which gas turbine engines operate.
  • Tungsten is also added for its solid solution strengthening properties. Tungsten is thought to be a more potent solid solution strengthener than either cobalt or iron, but cannot be added in large quantities (i.e. at amounts significantly above 5.5%) due to the increased promotion of deleterious intermetallics, and its adverse effect upon the alloy's density.
  • the amount of tungsten present in the composition is unusually high for a high temperature nickel based alloy, and has been found to greatly contribute to the high ultimate tensile strength of the alloy at high temperatures.
  • Niobium will partition to the gamma prime phase and provide strengthening, resisting the movement of dislocations through the gamma prime phase. However, the addition of too much niobium (i.e. significantly more than 1.5%) will result in the precipitation of the deleterious delta phase at the temperatures at which gas turbine components operate, which is not desired in the present invention. Niobium is also a potent metal carbide former (in conjunction with the carbon present in the alloy), which will improve dwell crack properties.
  • Tantalum is an effective gamma prime strengthener, preventing the movement of dislocations, which will give improved creep and other mechanical properties.
  • the addition of tantalum will significantly increase the density and elemental cost of the alloy, which are both undesirable.
  • tantalum can be added in greater quantities before the alloy stability is compromised (up to 1 %). Tantalum is also a potent metal carbide former.
  • Silicon promotes the formation of a more stable and resistive oxide scale than chromium alone. However, its introduction leads to great instability in anything other than small quantities (i.e more than 1.5%), which has a significant impact upon the mechanical properties of articles formed from the alloy. While a large quantity of silicon suppresses formation of the deleterious eta phase, too much promotes the formation of the deleterious G phase.
  • Molybdenum is optionally present in the alloy. Molybdenum is postulated to have a positive effect on the environmental resistance of the alloy. It will also act as a solid solution strengthener but is more prone to promoting the formation of deleterious phases then other solid solution strengthening elements.
  • Aluminium is essential for promoting the formation of the gamma prime phase, which provides the major strengthening mechanism for the alloy.
  • the control of its quantity is crucial to achieve the correct balance of properties (particularly in terms of the ratio of aluminium. Too much aluminium, and the alloy will be unprocessible (i.e., difficult to weld) - too little, and the alloy will have insufficient mechanical strength.
  • aluminium will improve the oxidation resistance and lower the density of the alloy, which are both highly important considerations in aerospace gas turbine engine components.
  • Titanium will strengthen the gamma prime phase as well as increase the fraction of gamma prime present. It will also reduce the density of the alloy. The addition of too much titanium (i.e. significantly more than 2.4%) will promote the formation of the deleterious eta phase, promote the formation of too much gamma prime, and may compromise the environmental resistance of the alloy by increasing oxide thickening rates.
  • the aluminium to titanium ratio is generally greater than 1:1. This, along with a Cr/Ti ratio greater than 10:1, promotes a good oxidation and corrosion resistance. However, it is still recognised the significant strengthening benefit that Ti adds and unlike most other strengthening elements, it has no density penalty.
  • Carbon, boron and zirconium are added in small amounts (as shown in table 2) to form carbides and borides on the grain boundaries, which strengthen the grain boundaries of the alloys.
  • Their amounts have been empirically optimised to prevent crack dwell fatigue and also prevent melt anomalies which will improve weldability.
  • the described alloy compositions can be used in various component manufacture methods, such as any of powder metallurgy methods, casting or laser deposition welding.
  • the described alloy compositions are particularly suitable for laser deposition welding, since the alloy is highly stable, and does not tend to microsegregate when melted.
  • This heat treatment may be performed either above or below the gamma prime solvus temperature to obtain the desired gamma prime precipitate distribution.
  • this may be followed by an ageing treatment, which nominally would be 4-16 hours at 850-900 °C.
  • composition 1 Various material properties of composition 1 are shown in Figs. 3 to 6 .
  • Fig. 3 shows, plotted against temperature, the rupture stress, ultimate tensile stress, yield stress and design space (i.e. the minimum of the rupture stress, ultimate tensile stress (UTS), and yield stress) of composition 1 (labelled as "optimised combustor” on the graph), compared to equivalent values of Nimonic 263TM.
  • the design space essentially represents the maximum stress that can be applied to the alloy prior to a failure of some sort at a given temperature.
  • composition 1 has a higher predicted yield stress and stress rupture behaviour than Nimonic 263TM, but a lower predicted UTS.
  • Fig. 4 compares the design space of composition 1 (again labelled as "optimised combustor alloy"), compared to Nimonic 263TM, Haynes 282, Haynes 230 and Haynes 188.
  • the design space of alloy composition 1 of the present disclosure is higher than all of the prior alloys at lower temperatures (less than 600°C), and only lower than Haynes 282 at higher temperatures.
  • Fig. 5 shows the relative molar fractions of various phases in composition 1, as plotted against temperature. As can be seen, at temperatures around 900°C, substantially only gamma and gamma prime are present. No precipitation of sigma ( ⁇ ) is observed, demonstrating that the desired level of microstructural stability has been achieved for use as an alloy that can be formed by direct laser deposition, and used as a combustor liner material.
  • Fig. 6 shows the activity of chromium of the alloy of composition 1, plotted as a function of temperature, in comparison to that of Nimonic 263. As can been seen, the chromium activity of the alloy of composition 1 is significantly higher, resulting in better environmental resistance (i.e. approximately half the oxidation rate of Nimonic 263).
  • Fig. 7 shows an image from a scanning electron microscope of the alloy of composition 1 after is has been subject to temperatures of 750°C for 1000 hours. As can be seen, there is substantially no microstructural segregation or cracking. Consequently, the alloy is resistant to high dwell temperatures.
  • this alloy is intended to be manufactured using direct laser deposition, however it is also suitable for other commonly used manufacturing techniques such as casting, powder processing and welding.
  • compositions which possess a number of advantageous properties compared to prior alloys.
  • main advantages of the new alloy in comparison to the prior art are:
  • alloys Although the description refers to the described alloys as being particularly useful for forming combustor linings for gas turbine engines, the alloys could also be used to form other components.
  • alloy is described as being subjected to heat treatment, other material processing methods could be used to generate articles from the alloy compositions having the required properties.

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  • Chemical & Material Sciences (AREA)
  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • Materials Engineering (AREA)
  • Metallurgy (AREA)
  • Organic Chemistry (AREA)
  • Combustion & Propulsion (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
  • Powder Metallurgy (AREA)

Claims (11)

  1. Alliage, constitué de, en pourcentage pondéral, entre 18 et 20 % de chrome, entre 8,5 % et 12 % de cobalt, entre 6,4 et 9,4 % de fer, jusqu'à 1,5 % de manganèse, entre 3,5 et 5,3 % de tungstène, jusqu'à 1,5 % de niobium, entre 0,4 % et 1 % de tantale, jusqu'à 1,5 % de silicone, entre 1,5 et 2,3 % d'aluminium, entre 0,5 et 2,4 % de titane, entre 0,005 et 0,04 % de carbone, entre 0,005 et 0,07 % de bore et entre 0,02 et 0,10 % de zirconium, facultativement, jusqu'à 0,5 % de molybdène, le reste étant constitué de nickel à l'exception des impuretés fortuites.
  2. Alliage selon la revendication 1, l'alliage étant constitué de, en pourcentage pondéral, entre 18 et 20 % de chrome, entre 8,5 % et 12 % de cobalt, entre 6,4 et 9,4 % de fer, jusqu'à 1,5 % de manganèse, entre 3,5 et 5,3 % de tungstène, jusqu'à 1,5 % de niobium, entre 0,4 % et 1 % de tantale, entre 0,1 % et 1,5 % de silicone, entre 1,5 et 2,3 % d'aluminium, entre 1 et 2,4 % de titane, entre 0,005 et 0,04 % de carbone, entre 0,005 et 0,07 % de bore et entre 0,02 et 0,10 % de zirconium, le reste étant constitué de nickel à l'exception des impuretés fortuites.
  3. Alliage selon la revendication 1, l'alliage comprenant entre 0,1 % et 1,5 % en poids de silicone.
  4. Composition d'alliage selon la revendication 1 ou la revendication 2, l'alliage étant constitué de, en pourcentage pondéral, 20 % de chrome, entre 9,3 % et 11,3 % de cobalt, entre 7,4 et 8,4 % de fer, entre 0,2 et 0,4 % de manganèse, entre 4,2 et 4,6 % de tungstène, entre 1 et 1,2 % de niobium, entre 0,6 % et 0,8 % de tantale, entre 0,2 et 0,4 % de silicone, entre 1,8 et 2,0 % d'aluminium, entre 1,6 et 1,8 % de titane, entre 0,015 et 0,025 % de carbone, entre 0,015 et 0,025 % de bore et entre 0,05 et 0,07 % de zirconium, le reste étant constitué de nickel à l'exception des impuretés fortuites.
  5. Composition d'alliage selon l'une quelconque des revendications précédentes, l'alliage étant constitué de, en pourcentage pondéral, 20 % de chrome, 10,3 % de cobalt, 7,9 % de fer, 0,3 % de manganèse, 4,4 % de tungstène, 1,1 % de niobium, entre 0,7 % de tantale, 0,3 % de silicone, 1,9 % d'aluminium, 1,7 % de titane, entre 0,015 et 0,025 % de carbone, entre 0,015 et 0,025 % de bore et entre 0,05 et 0,07 % de zirconium, le reste étant constitué de nickel à l'exception des impuretés fortuites.
  6. Alliage selon l'une quelconque des revendications précédentes, dans lequel les impuretés fortuites sont notamment de l'oxygène.
  7. Alliage selon la revendication 1 ou la revendication 2, dans lequel le rapport en poids du titane à l'aluminium est inférieur à 0,5.
  8. Alliage selon l'une quelconque des revendications précédentes, dans lequel le rapport en poids du chrome au titane est supérieur à 10.
  9. Composant de turbine à gaz constitué d'un alliage selon l'une quelconque des revendications précédentes.
  10. Composant de turbine à gaz selon la revendication 9, le composant comprenant une chemise thermique.
  11. Turbine à gaz comprenant un composant selon la revendication 9 ou la revendication 10.
EP15163869.9A 2014-05-14 2015-04-16 Composition d'alliage de nickel Not-in-force EP2944704B1 (fr)

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US10221468B2 (en) 2016-06-30 2019-03-05 General Electric Company Article and additive manufacturing method for making
US10577679B1 (en) 2018-12-04 2020-03-03 General Electric Company Gamma prime strengthened nickel superalloy for additive manufacturing

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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPS53108022A (en) * 1977-03-04 1978-09-20 Hitachi Ltd Iron-nickel-chromium-molybdenum alloy of high ductility
US5480283A (en) * 1991-10-24 1996-01-02 Hitachi, Ltd. Gas turbine and gas turbine nozzle
KR100372482B1 (ko) * 1999-06-30 2003-02-17 스미토모 긴조쿠 고교 가부시키가이샤 니켈 베이스 내열합금
JP3596430B2 (ja) * 1999-06-30 2004-12-02 住友金属工業株式会社 Ｎｉ基耐熱合金
JP4277113B2 (ja) * 2002-02-27 2009-06-10 大同特殊鋼株式会社 耐熱ばね用Ｎｉ基合金
US8066938B2 (en) * 2004-09-03 2011-11-29 Haynes International, Inc. Ni-Cr-Co alloy for advanced gas turbine engines

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EP2944704A1 (fr) 2015-11-18
US20150329941A1 (en) 2015-11-19

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