Alloy, protective layer, and component

Abstract

Known protective layers with a high Cr content and additionally silicon form brittle phases that become even more brittle under the influence of carbon during use. The protective layer according to the invention has the composition NiCoCrAlYTaSiHf: 22% - 24% cobalt (Co), 14% - 20% chromium (Cr), 8.0% - 12.0% aluminium (Al), 0.2% - 0.6% at least one metal from the group comprising scandium (Sc) and / or rare-earth elements, 1.0% to 3.0% tantalum (Ta), 0.1% to 0.7% silicon (Si), 0.1% to 1.0% hafnium (Hf), nickel (Ni).

Description

Description TITLE Alloy, protective layer and component TECHNICAL FIELD [1] The invention relates to an alloy according to claim 1, a protective layer for protecting a component against corrosion and / or oxidation, in particular at high temperatures according to claim 9 and a component according to claim 11. BACKGROUND D [2] A large number of protective coatings for metallic components intended to increase their corrosion and / or oxidation resistance are known in the art. Most of these protective coatings are known collectively under the name MCrAlY, where M stands for at least one of the elements from the group comprising iron, cobalt, and nickel, and other essential components are chromium, aluminum, and yttrium. [3] Typical coatings of this type are known from US patents 4,005,989 and 4,034,142. [4] The effort to increase the inlet temperatures in both stationary gas turbines and aircraft engines is of great importance in the field of gas turbines, as the inlet temperatures are important determinants of the thermodynamic efficiencies achievable with gas turbines. Through the use of specially developed alloys as base materials for components subject to high thermal stress, such as guide vanes and rotor blades, in particular through the use of single-crystal superalloys, inlet temperatures of well over 1273 K are possible. Meanwhile, the state of the art allows inlet temperatures of 1223K and more in stationary gas turbines and 1373K and more in gas turbines of aircraft engines. [5] Examples of the construction of a turbine blade with a single-crystal substrate, which in turn can have a complex structure, can be found in WO 91 / 01433 A1. [6] While the physical resilience of the base materials now developed for the highly stressed components is largely unproblematic with regard to possible further increases in inlet temperatures, protective coatings must be used to achieve sufficient resistance to oxidation and corrosion. In addition to the sufficient chemical resistance of a protective coating under the attacks expected from flue gases at temperatures in the order of 1273 K, a protective coating must also have sufficiently good mechanical properties, not least with regard to the mechanical interaction between the protective coating and the base material. In particular, the protective coating must be sufficiently ductile to be able to follow any deformation of the base material and not crack, as this would create points of attack for oxidation and corrosion.The typical problem here is that an increase in the proportions of elements such as aluminum and chromium, which can improve the resistance of a protective layer to oxidation and corrosion, leads to a deterioration in the ductility of the protective layer, so that mechanical failure, in particular the formation of cracks, can be expected under the mechanical load typically encountered in a gas turbine. SUMMARY OF THE INVENTION [7] Accordingly, the object of the invention is to provide an alloy and a protective layer which has good high-temperature resistance to corrosion and oxidation, has good long-term stability and which is also particularly well adapted to mechanical stress which is to be expected in particular in a gas turbine at a high temperature. [8] The object is achieved by an alloy according to claim 1 and a protective layer according to claim 9. [9] A further object of the invention is to provide a component which has increased protection against corrosion and oxidation.

[0010] The object is also achieved by a component according to claim 11, in particular a component of a gas turbine or steam turbine, which has a protective layer of the type described above for protection against corrosion and oxidation at high temperatures.

[0011] The subclaims list further advantageous measures which can be combined with each other in any advantageous manner.

[0012] The invention is explained in more detail below.

[0013] The figures and the description represent only embodiments of the invention.

[0014] It shows Figure 1 a layer system with a protective layer, Figure 2 Compositions of superalloys, Figure 3 a gas turbine, Figure 4 a turbine blade and Figure 5 a combustion chamber. Description of the characters

[0015] According to the invention, a protective layer 7 (Fig. 1) for protecting a component against corrosion and oxidation at a high temperature essentially comprises the following elements (proportions given in wt.%): nickel Cobalt (Co): 22.0% - 24.0%, in particular 23.0%, Chromium (Cr): 14.0% - 20%, in particular 14.0% - 16.0%, Aluminum (AI): 8.0% - 12.0%, in particular 9.0% - 11.0%, 0.2% - 0.6% rare earth element (yttrium (Y), ...) and / or scandium (Sc), Tantalum (Ta): 1.0% - 3.0%, in particular 1.4% - 2.0% tantalum, Hafnium (Hf): 0.1% - 1.0%, in particular 0.2% - 0.8% hafnium, Silicon (Si): 0.1% - 0.7%, in particular 0.2% - 0.5% silicon.

[0016] The list of alloying elements Ni, Co, Cr, AI, Y, Ta, Si, Hf is preferably not exhaustive.

[0017] Nickel preferentially forms the matrix.

[0018] Preferably, the list of Ni, Co, Cr, AI, Y, Ta, Si, Hf is exhaustive.

[0019] The contents of the alloying elements Co, Cr, AI, Y, Ta, Si, Hf have the following advantages: Medium-high Co content: broadening of the beta / gamma field, avoiding brittle phases such as the alpha phases. Medium Cr content: Sufficiently high to increase the activity of AI for AI2O3 formation; low enough to avoid brittle phases (alpha chromium or sigma phase). Medium-high Al content: Sufficiently high for Al activity to form a stable AlO layer; low enough to avoid embrittlement effects. Low Y content: Sufficiently high to produce enough Y-aluminate to form Y-containing pegs even with low oxygen contamination; low enough to negatively accelerate the oxide growth of the ALOs layer. Tantalum has a positive influence on the phase stability of the y'-phase or shifts the transition to higher temperatures and thus slows down the phase degradation through the consumption of aluminum in the layer. Silicon and hafnium extend the lifetime and adhesion of the layer.

[0020] Good results are achieved with Ni-23Co-15Cr-8.2AI-0.4Y-1.5Ta- 0.3Hf-0.2Si or Ni-22Co-16Co-9.5AI-0.4Y-1 .2Ta-0.2Hf-0.15Si

[0021] It should be noted that the proportions of the individual elements are specifically calibrated with regard to their effects, which are particularly important in relation to the element silicon. If the proportions are such that no silicon precipitation occurs, brittle phases are advantageously prevented from developing during the protective coating's use, thus improving and extending its service life.

[0022] This is achieved not only by a low chromium content, but also by precisely measuring the aluminum content, taking into account the influence of aluminum on phase formation.

[0023] In interaction with the reduction of the brittle phases, which have a negative effect especially under higher mechanical properties, the mechanical properties are improved by reducing the mechanical stresses through the selected nickel content.

[0024] The protective coating exhibits good corrosion resistance, particularly good oxidation resistance, and is also characterized by particularly good ductility properties, making it particularly suitable for use in a gas turbine 100 (Fig. 3) with a further increase in inlet temperature. Embrittlement hardly occurs during operation, as the coating contains hardly any chromium-silicon precipitates, which become brittle over time.

[0025] The powders are applied, for example, by plasma spraying (APS, LPPS, VPS, etc.) to form a protective layer and thus exhibit a characteristic microstructure. Other processes are also conceivable (PVD, CVD, SPPS, etc.).

[0026] The described protective layer 7 also acts as an adhesion promoter layer to a superalloy.

[0027] Further layers, in particular ceramic thermal insulation layers 10, can be applied to this protective layer 7.

[0028] In a component 1, the protective layer 7 is advantageously applied to a substrate 4 made of a nickel- or cobalt-based superalloy (Figure 2).

[0029] Compositions of this type are known as casting alloys under the designations GTD222, IN939, IN6203 and Udirnet 500.

[0030] Further alternatives for the substrate 4 (Figure 2) of the component 1, 120, 130, 155 are listed in Figure 2.

[0031] The thickness of the protective layer 7 on the component 1 is preferably set to a value between approximately 100pm and 500pm.

[0032] The protective layer 7 is particularly suitable for protecting the component 1, 120, 130, 155 against corrosion and oxidation while the component is exposed to a flue gas at a material temperature of approximately 950°C, or approximately 1100°C in the case of aircraft turbines.

[0033] The protective layer 7 according to the invention is thus particularly qualified for protecting a component of a gas turbine 100, in particular a guide vane 120, rotor blade 130 or a heat shield element 155, which is exposed to hot gas in front of or in the turbine of the gas turbine 100 or the steam turbine.

[0034] The protective layer 7 can be used as a coating layer (protective layer is the outermost layer) or as an adhesion promoter layer (protective layer is an intermediate layer).

[0035] Figure 1 shows a layer system 1 as a component.

[0036] The layer system 1 has a substrate 4.

[0037] The substrate 4 can be metallic and / or ceramic. Particularly in turbine components, such as turbine rotor blades 120 (Fig. 4) or guide vanes 130 (Figs. 3, 4), heat shield elements 155 (Fig. 5), and other casing parts of a steam or gas turbine 100 (Fig. 3), the substrate 4 comprises, and in particular consists of, a nickel-, cobalt-, or iron-based superalloy.

[0038] Preferably, nickel-based superalloys (Fig. 2) are used.

[0039] The protective layer 7 according to the invention is present on the substrate 4.

[0040] Preferably, this protective layer 7 is applied by plasma spraying (VPS, LPPS, APS, ...).

[0041] This can be used as an outer layer (not shown) or intermediate layer (Fig. 1 ).

[0042] Preferably, a ceramic thermal insulation layer 10 is present on the protective layer 7.

[0043] Preferably, the layer system consists of substrate 4, protective layer 7 and ceramic thermal insulation layer 10, optionally a TGO under the thermal insulation layer 10.

[0044] The protective layer 7 can be applied to newly manufactured components and remanufactured components from refurbishment.

[0045] Refurbishment means that components 1 are separated from their layers (thermal insulation layer) after use, if necessary, and corrosion and oxidation products are removed, for example, through acid treatment (acid stripping). Cracks may still need to be repaired. Afterward, such a component can be recoated, as the substrate 4 is very expensive.

[0046] Figure 3 shows an example of a gas turbine 100 in a longitudinal section.

[0047] The gas turbine 100 has inside a rotor 103 with a shaft 101, which is rotatably mounted about a rotation axis 102 and is also referred to as a turbine rotor.

[0048] Along the rotor 103, there follow an intake housing 104, a compressor 105, a toroidal combustion chamber 110, in particular an annular combustion chamber, with several coaxially arranged burners 107, a turbine 108 and the exhaust housing 109.

[0049] The annular combustion chamber 110 communicates with a hot gas duct 111, for example an annular one. There, for example, four turbine stages 112 connected in series form the turbine 108.

[0050] Each turbine stage 112 is formed, for example, from two blade rings. Viewed in the flow direction of a working medium 113, a row of guide vanes 115 is followed in the hot gas duct 111 by a row 125 formed from rotor blades 120.

[0051] The guide vanes 130 are attached to an inner housing 138 of a stator 143, whereas the rotor blades 120 of a row 125 are attached to the rotor 103, for example by means of a turbine disk 133.

[0052] A generator or a working machine (not shown) is coupled to the rotor 103.

[0053] During operation of the gas turbine 100, air 135 is drawn in by the compressor 105 through the intake casing 104 and compressed. The compressed air provided at the turbine end of the compressor 105 is fed to the burners 107, where it is mixed with a fuel. The mixture is then combusted in the combustion chamber 110 to form the working medium 113. From there, the working medium 113 flows along the hot gas duct 111 past the guide vanes 130 and the rotor blades 120. At the rotor blades 120, the working medium 113 expands, transmitting momentum, so that the rotor blades 120 drive the rotor 103, which drives the driven machine coupled to it.

[0054] The components exposed to the hot working medium 113 are subject to thermal stress during operation of the gas turbine 100. The guide vanes 130 and rotor blades 120 of the first turbine stage 112, viewed in the flow direction of the working medium 113, are subjected to the greatest thermal stress, along with the heat shield elements lining the annular combustion chamber 110. To withstand the temperatures prevailing there, these elements can be cooled using a coolant.

[0055] Likewise, substrates of the components can have a directed structure, i.e. they are single-crystalline (SX structure) or have only longitudinally directed grains (DS structure).

[0056] Iron-, nickel-, or cobalt-based superalloys, for example, are used as materials for the components, in particular for the turbine blades 120, 130 and components of the combustion chamber 110. Such superalloys are known, for example, from EP 1 204 776 B1, EP 1 306 454, EP 1 319 729 A1, WO 99 / 67435, or WO 00 / 44949.

[0057] The guide vane 130 has a guide vane root (not shown here) facing the inner casing 138 of the turbine 108 and a guide vane tip opposite the guide vane root. The guide vane tip faces the rotor 103 and is secured to a mounting ring 140 of the stator 143.

[0058] Figure 4 shows a perspective view of a rotor blade 120 or guide vane 130 of a turbomachine extending along a longitudinal axis 121. The turbomachine can be a gas turbine of an aircraft or a power plant for generating electricity, a steam turbine, or a compressor. The blade 120, 130 has, along the longitudinal axis 121, a mounting region 400, an adjacent blade platform 403, a blade airfoil 406, and a blade tip 415. As a guide vane 130, the blade 130 can have a further platform at its blade tip 415 (not shown). A blade root 183 is formed in the mounting region 400, which serves to attach the rotor blades 120, 130 to a shaft or a disk (not shown). The blade root 183 is designed, for example, as a hammer head. Other designs such as a fir tree root or dovetail root are possible. The blade 120, 130 has a leading edge 409 and a trailing edge 412 for a medium flowing past the blade 406.In conventional blades 120, 130, solid metallic materials, in particular superalloys, are used in all areas 400, 403, 406 of the blade 120, 130. Such superalloys are known, for example, from EP 1 204 776 B1, EP 1 306454, EP 1 319 729 A1, WO 99 / 67435 or WO 00 / 44949. The blade 120, 130 can be manufactured by a casting process, also by means of directional solidification, by a forging process, by a milling process or combinations thereof. Workpieces with a single-crystal structure or structures are used as components for machines that are exposed to high mechanical, thermal and / or chemical loads during operation. The production of such single-crystal workpieces takes place, for example, by directional solidification from the melt. These are casting processes in which the liquid metallic alloy is cast into a single-crystalline structure, ieto a single-crystal workpiece, or directionally solidified. In this process, dendritic crystals are aligned along the heat flow and form either a columnar grain structure (i.e., grains that run the entire length of the workpiece) or a columnar grain structure (i.e., grains that run the entire length of the workpiece). be) or a single-crystalline structure, i.e. the entire workpiece consists of a single crystal. In these processes, the transition to globulitic (polycrystalline) solidification must be avoided, since undirected growth necessarily leads to the formation of transverse and longitudinal grain boundaries, which destroy the good properties of the directionally solidified or single-crystalline component. When we generally talk about directionally solidified structures, we mean both single crystals that have no grain boundaries or at most small-angle grain boundaries, and columnar crystal structures that have grain boundaries running longitudinally but no transverse grain boundaries. These latter crystalline structures are also referred to as directionally solidified structures. Such processes are known from US Pat. No. 6,024,792 and EP 0 892 090 A1.

[0059] Likewise, the blades 120, 130 can have protective coatings 7 according to the invention against corrosion or oxidation. The density is preferably 95% of the theoretical density. A protective aluminum oxide layer (TGO = thermal grown oxide layer) forms on the MCrAIX layer (as an intermediate layer or as the outermost layer). A thermal insulation layer can also be present on the MCrAIX, which is preferably the outermost layer and consists, for example, of ZrO2, Y2O3-ZrO2, i.e., it is not, partially, or fully stabilized by yttrium oxide and / or calcium oxide and / or magnesium oxide.

[0060] The thermal barrier coating covers the entire MCrAIX layer. Suitable coating processes, such as electron beam evaporation (EB-PVD), create columnar grains in the thermal barrier coating. Other coating processes are conceivable, such as atmospheric plasma spraying (APS), LPPS, VPS, or CVD. The thermal barrier coating can contain porous grains with micro- or macrocracked surfaces for improved thermal shock resistance. Therefore, the thermal barrier coating is preferably more porous than the MCrAIX layer.

[0061] The blade 120, 130 can be hollow or solid. If the blade 120, 130 is to be cooled, it is hollow and may also have film cooling holes 418 (indicated by dashed lines).

[0062] Figure 5 shows a combustion chamber 110 of the gas turbine 100. The combustion chamber 110 is designed, for example, as a so-called annular combustion chamber, in which a plurality of burners 107 arranged in the circumferential direction around a rotation axis 102 open into a common combustion chamber space 154, which generate flames 156. For this purpose, the combustion chamber 110 is designed as a whole as an annular structure that is positioned around the rotation axis 102. To achieve a comparatively high level of efficiency, the combustion chamber 110 is designed for a comparatively high temperature of the working medium M of approximately 1000°C to 1600°C. To enable a comparatively long operating time even with these operating parameters, which are unfavorable for the materials, the combustion chamber wall 153 is provided on its side facing the working medium M with an inner lining formed from heat shield elements 155. Due to the high temperatures inside the combustion chamber 110, a heat shield element 155 orA cooling system may be provided for their holding elements. The heat shield elements 155 are then hollow, for example, and may also have cooling holes (not shown) opening into the combustion chamber 154. Each heat shield element 155 made of an alloy is equipped with a particularly heat-resistant protective layer (MCrAIX layer and / or ceramic coating) on ​​the working medium side or is made of high-temperature-resistant material (solid ceramic bricks). These protective layers 7 can be similar to those used on turbine blades. A ceramic thermal barrier coating, for example, may also be present on the MCrAIX and consists, for example, of ZrO2, Y2O3-ZrO2, i.e. it is not, partially, or fully stabilized by yttrium oxide and / or calcium oxide and / or magnesium oxide. Rod-shaped grains are created in the thermal barrier coating using suitable coating processes such as electron beam evaporation (EB-PVD).Other coating processes are conceivable, e.g. Atmospheric plasma spraying (APS), LPPS, VPS, or CVD. The thermal barrier coating can contain porous, micro- or macro-cracked grains for improved thermal shock resistance.

[0063] Refurbishment means that turbine blades 120, 130, heat shield elements 155 are, if necessary, Protective layers must be removed (e.g., by sandblasting). This is followed by the removal of corrosion and / or oxidation layers or products. If necessary, cracks in the turbine blades 120, 130, or the heat shield element 155 are also repaired. The turbine blades 120, 130, and the heat shield elements 155 are then recoated and reused.

Claims

Patent claims 1. Alloy containing at least the following elements (in wt.%): 22.0% - 24.0% cobalt (Co), in particular 23.0% cobalt (Co), 14.0% - 20.0% chromium (Cr), in particular 14.0% - 16.0% chromium (Cr), 8.0% - 12.0% aluminum (AI), in particular 9.0% - 11.0% aluminum (AI), 0.2% - 0.6%, of at least one metal from the group comprising scandium (Sc) and / or the rare earth elements, in particular yttrium (Y), 1.0% to 3.0% tantalum (Ta) 0.1% to 0.7% silicon (Si), 0.1% to 1.0% hafnium (Hf), in particular 0.2% to 0.8% hafnium (Hf), Nickel (Ni), especially residual nickel (Ni).

2. Alloy according to claim 1, containing 0.4 wt.% yttrium (Y), in particular only yttrium (Y) containing from the group.

3. Alloy according to claim 1 or 2, not containing rhenium (Re) or ruthenium (Ru).

4. Alloy according to one or more of the preceding claims, which contains 0.2% to 0.5% silicon (Si).

5. Alloy according to one or more of the preceding claims, which contains tantalum (Ta), in particular at least 1.4 wt.% tantalum (Ta), most particularly at most 2.0 wt.% tantalum (Ta).

6. Alloy according to one or more of the preceding claims, not containing zirconium (Zr) and / or not containing titanium (Ti) and / or not containing gallium (Ga) and / or not containing germanium (Ge) and / or not containing platinum (Pt) and / or not containing cerium (Ce) and / or not containing iron (Fe) and / or not containing palladium (Pd) and / or not containing boron (B) and / or not containing carbon (C).

7. Alloy according to one or more of the preceding claims, consisting of cobalt, chromium, aluminum, yttrium, tantalum, silicon, hafnium, nickel.

8. Alloy according to one or more of the preceding claims, in which nickel (Ni) forms the matrix.

9. Protective layer (7) for protecting a component (1) against corrosion and / or oxidation, in particular at high temperatures, which has the composition of the alloy according to one or more of claims 1 to 8.

10. A protective layer (7) according to claim 9, which is applied by plasma spraying, in particular APS, or high-velocity spraying (HVOF).

11. A component, in particular a component (120, 130, 155) of a gas turbine (100), in particular in which a substrate (4) of the component (120, 130, 155) is nickel-based or cobalt-based, which has a protective layer (7) according to claim 9 or 10 for protection against corrosion and oxidation at high temperatures, in which in particular a ceramic thermal barrier coating (10) is applied to the protective layer (7).

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