EP2354260A1 - Alloy, protective layer and component - Google Patents

Abstract

Alloy comprises (in wt.%) cobalt (18-22 (preferably 19-21)), chromium (22-26 (preferably 23-25)), aluminum (6-8 (preferably 6.5-7.5)), at least one element comprising scandium or rare earth elements, preferably yttrium (0.5-0.7), and nickel (balance). Independent claims are included for: (1) a protective layer (7) for protecting a component against corrosion and/or oxidation, preferably at high temperatures, comprising the composition of the alloy, and is present as a single layer; and (2) the component, preferably a component of a gas turbine, preferably in which a substrate (4) of the component is nickel-based or cobalt-based, comprising the protective layer, preferably a metallic protective layer, in which a ceramic thermal barrier layer (10) is applied on the protective layer.

Description

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 10 and a component according to claim 11.

Protective layers for metallic components intended to increase their corrosion resistance and / or oxidation resistance are known in the art in large numbers. Most of these protective layers are known under the collective name MCrAlY, where M represents at least one of the elements selected from the group consisting of iron, cobalt and nickel and further essential components are chromium, aluminum and yttrium.

Typical coatings of this kind are from the U.S. Patents 4,005,989 and 4,034,142 known.

Also known is the addition of rhenium (Re) to NiCoCrAlY alloys.

Efforts to increase the inlet temperatures of both stationary gas turbines and aircraft engines are of great importance in the gas turbine art because inlet temperatures are important determinants of thermodynamic efficiencies achievable with gas turbines. Due to the use of specially developed alloys as base materials for components that are subject to high thermal loads such as guide vanes and rotor blades, in particular through the use of monocrystalline superalloys, inlet temperatures of well over 1000 ° C are possible. Meanwhile, the prior art allows inlet temperatures of 950 ° C and more in stationary gas turbines and 1100 ° C and more in gas turbines of aircraft engines.

Examples of the construction of a turbine blade with a monocrystalline substrate, which in turn can be complex, emerge from the WO 91/01433 Al ,

While the physical strength of the now developed base materials for the highly loaded components with regard to possible further increases in the inlet temperatures is largely unproblematic, must be used to achieve a sufficient resistance to oxidation and corrosion on protective layers. In addition to the sufficient chemical resistance of a protective layer under the attacks that are expected of flue gases at temperatures in the order of 1000 ° C, a protective layer must also have sufficient mechanical properties, not least in view of the mechanical interaction between the protective layer and the base material , to have. In particular, the protective layer must be sufficiently ductile in order to be able to follow any deformations of the base material and not to break, since in this way points of attack for oxidation and corrosion would be created.

Accordingly, it is an object of the present invention to provide an alloy and a protective layer which has good high-temperature resistance in corrosion and oxidation, has good long-term stability and, in addition, a mechanical stress to be expected particularly in a gas turbine at a high temperature well adjusted.

The object is achieved by an alloy according to claim 1 and a protective layer according to claim 10.

Another object of the invention is to provide a component which has increased protection against corrosion and oxidation.

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

In the subclaims further advantageous measures are listed, which can be linked to each other in an advantageous manner.

The invention is u. a. based on the knowledge that the protective layer in the layer and in the transition region between protective layer and base material brittle rhenium precipitates shows. During operation, these brittle phases, which form increasingly with time and temperature, lead to pronounced longitudinal cracks in the layer as well as in the interface layer base material with subsequent detachment of the layer. In addition, the brittleness of the rhenium precipitates increases as a result of the interaction with carbon, which can diffuse into the layer from the base material or diffuse into the layer during a heat treatment in the furnace through the surface. Oxidation of the rhenium phases further enhances the driving force for crack formation.

The invention will be explained in more detail below.

Figur 1

ein Schichtsystem mit einer Schutzschicht,

Figur 2

Zusammensetzungen von Superlegierungen,

Figur 3

eine Gasturbine,

Figur 4

eine Turbinenschaufel und

Figur 5

eine Brennkammer.

Show it

FIG. 1

a layer system with a protective layer,

FIG. 2

Compositions of superalloys,

FIG. 3

a gas turbine,

FIG. 4

a turbine blade and

FIG. 5

a combustion chamber.

The figures and the description represent only embodiments of the invention.

• 18% bis 22% Kobalt (Co),
• 6% bis 8% Aluminium (Al),
• 0,5% bis 0,7% Yttrium (Y) und/oder zumindest ein äquivalentes Metall aus der Gruppe umfassend Scandium und die Elemente der Seltenen Erden,
• 22% bis 20% Chrom (Cr),
• Nickel (Ni) (NiCoCrAlY).

According to the invention, a protective layer 7 (FIG. Fig. 1 ) to protect a component against corrosion and oxidation at a high temperature substantially the following elements (see the shares in% by weight):

• 18% to 22% cobalt (Co),
• 6% to 8% aluminum (Al),
• 0.5% to 0.7% yttrium (Y) and / or at least one equivalent metal from the group comprising scandium and the elements of the rare earths,
• 22% to 20% chromium (Cr),
• Nickel (Ni) (NiCoCrAlY).

This listing is not exhaustive.

An advantageous embodiment consists of the elements nickel, cobalt, chromium, aluminum and yttrium

At higher oxidation load (pure combustion gas), more oxygen must be bound by yttrium, so that the protective aluminum oxide layer can not grow too fast, and then the yttrium value is advantageously up to 0.7wt%. However, the content of yttrium should generally not be too high in the alloy, otherwise this leads to embrittlement.

• Ni - 20Co - 24Cr - 7A1 - 0,6Y .

A preferred embodiment is:

• Ni - 20Co - 24Cr - 7A1 - 0,6Y.

It should be noted that the proportions of the individual elements are specially tuned with regard to their effects, which are to be seen in particular in connection with the non-existent element rhenium. If the proportions are so large, can be dispensed with the addition of rhenium (Re), so that no rhenium precipitates form. There are advantageously no brittle phases during the use of the protective layer, so that the runtime behavior is improved and extended.

In interaction with the reduction of the brittle phases, which have a negative effect especially under higher mechanical properties, by reducing the mechanical Stresses due to the selected nickel content improves the mechanical properties.

The protective layer has, with good corrosion resistance, a particularly good resistance to oxidation and is also distinguished by particularly good ductility properties, so that it is particularly qualified for use in a gas turbine 100 (FIG. Fig. 3 ) with a further increase in the inlet temperature.

The powders are applied for example by plasma spraying (APS, LPPS, VPS, ...). Other methods are also conceivable (PVD, CVD, cold gas spraying, ...).

The protective layer 7 described also acts as a primer layer to a superalloy.

Preferably, only a single protective layer 7 is used for the component, ie no duplex layer for the bondcoat. On this protective layer 7 more layers, in particular ceramic thermal barrier coatings 10 can be applied.

In the case of a component 1, the protective layer 7 is advantageously applied to a substrate 4 made of a nickel or cobalt-based superalloy.

In particular the following composition is suitable as substrate (data in wt%): 0.1% up to 0.15% carbon 18% up to 22% chrome 18% until 19% cobalt 0% up to 2% tungsten 0% to 4% molybdenum 0% up to 1.5% tantalum 0% up to 1% niobium 1% to 3% aluminum 2% to 4% titanium 0% up to 0.75% hafnium optionally small amounts of boron and / or zirconium, balance nickel.

Compositions of this type are known as casting alloys under the designations GTD222, IN939, IN6203 and Udimet 500.

Further alternatives for the substrate 4 of the component 1, 120, 130, 155 are in the FIG. 2 listed.

The thickness of the protective layer 7 on the component 1 is preferably dimensioned to a value of between about 100 μm and 300 μm.

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 about 950 ° C, in aircraft turbines also at about 1100 ° C.

The protective layer 7 according to the invention is therefore particularly qualified for protecting a component of a gas turbine 100, in particular a guide blade 120, blade 130 or a heat shield element 155, which is acted upon by hot gas before or in the turbine of the gas turbine 100 or the steam turbine.

The protective layer 7 can be used as an overlay (protective layer is the outer layer or as a bondcoat (protective layer is an intermediate layer).

It is preferably used as a "single" layer, i. there is no other metallic layer.

FIG. 1 shows a layer system 1 as a component.

The layer system 1 consists of a substrate 4.

The substrate 4 may be metallic and / or ceramic. Especially in turbine components, such as turbine run 120 ( Fig. 4 ) or vanes 130 (FIG. Fig. 3, 4 ), Heat shield elements 155 ( Fig. 5 ) and other housing parts of a steam or gas turbine 100 ( Fig. 3 ), the substrate 4 consists of a nickel-, cobalt- or iron-based superalloy. Preferably, nickel-based superalloys are used.

On the substrate 4, the protective layer 7 according to the invention is present. It is preferably used as a "single" layer, i. there is no other metallic layer. Preferably, this protective layer 7 is applied by plasma spraying (VPS, LPPS, APS1,...).

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

In the latter case, a ceramic thermal barrier coating 10 is present on the protective layer 7.

The protective layer 7 can be applied to newly manufactured components and refurbished components from the refurbishment.

Refurbishment means that after use, components 1 may be separated from layers (thermal barrier coating) and corrosion and oxidation products may be removed, for example by acid treatment (acid stripping). If necessary, cracks still have to be repaired. Thereafter, such a component can be coated again because the substrate 4 is very expensive.

The FIG. 3 shows by way of example a gas turbine 100 in a longitudinal partial section.

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

Along the rotor 103 follow one another an intake housing 104, a compressor 105, for example, a toroidal combustion chamber 110, in particular annular combustion chamber, with a plurality of coaxially arranged burners 107, a turbine 108 and the exhaust housing 109th

The annular combustion chamber 110 communicates with an annular annular hot gas channel 111, for example. There, for example, four turbine stages 112 connected in series form the turbine 108.

Each turbine stage 112 is formed, for example, from two blade rings. As seen in the direction of flow of a working medium 113, in the hot gas channel 111 of a row of guide vanes 115, a series 125 formed of rotor blades 120 follows.

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

Coupled to the rotor 103 is a generator or work machine (not shown).

During operation of the gas turbine 100, air 105 is sucked in and compressed by the compressor 105 through the intake housing 104. The compressed air provided at the turbine-side end of the compressor 105 is supplied to the burners 107 where it is mixed with a fuel. The mixture is then burned to form the working fluid 113 in the combustion chamber 110. From there, the working medium 113 flows along the hot gas channel 111 past the guide vanes 130 and the rotor blades 120. On the rotor blades 120, the working medium 113 expands in a pulse-transmitting manner so that the rotor blades 120 drive the rotor 103 and drive the machine coupled to it.

The components exposed to the hot working medium 113 are subject to thermal loads 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 in addition to the heat shield elements lining the annular combustion chamber 110.

To withstand the prevailing temperatures, they can be cooled by means of a coolant.

Likewise, substrates of the components can have a directional structure, ie they are monocrystalline (SX structure) or have only longitudinal grains (DS structure).

As the material for the components, in particular for the turbine blade 120, 130 and components of the combustion chamber 110, for example, iron-, nickel- or cobalt-based superalloys are used.

Such superalloys are for example from EP 1 204 776 B1 . EP 1 306 454 . EP 1 319 729 A1 . WO 99/67435 or WO 00/44949 known.

The vane 130 has a guide vane foot (not shown here) facing the inner casing 138 of the turbine 108 and a vane head opposite the vane root. The vane head faces the rotor 103 and fixed to a mounting ring 140 of the stator 143.

The FIG. 4 shows a perspective view of a blade 120 or guide vane 130 of a turbomachine, which extends along a longitudinal axis 121.

The turbomachine may be a gas turbine of an aircraft or a power plant for power generation, a steam turbine or a compressor.

The blade 120, 130 has along the longitudinal axis 121 consecutively a fastening region 400, a blade platform 403 adjacent thereto and an airfoil 406 and a blade tip 415.

As a guide blade 130, the blade 130 may have at its blade tip 415 another platform (not shown).

In the mounting region 400, a blade root 183 is formed, which serves for attachment of the blades 120, 130 to a shaft or a disc (not shown).

The blade root 183 is designed, for example, as a hammer head. Other designs as Christmas tree or Schwalbenschwanzfuß are possible.

The blade 120, 130 has a leading edge 409 and a trailing edge 412 for a medium flowing past the airfoil 406.

In conventional blades 120, 130, for example, solid metallic materials, in particular superalloys, are used in all regions 400, 403, 406 of the blade 120, 130.

Such superalloys are for example from EP 1 204 776 B1 . EP 1 306 454 . EP 1 319 729 A1 . WO 99/67435 or WO 00/44949 known.

The blade 120, 130 can be made by a casting process, also by directional solidification, by a forging process, by a milling process or combinations thereof.

Workpieces with a monocrystalline structure or structures are used as components for machines which are exposed to high mechanical, thermal and / or chemical stresses during operation.

The manufacture of such monocrystalline workpieces takes place, for example, by directed solidification from the melt. These are casting processes in which the liquid metallic alloy is transformed into a monocrystalline structure, i. to the single-crystal workpiece, or directionally solidified.

Here, dendritic crystals are aligned along the heat flow and form either a columnar grain structure (columnar, ie grains that run the entire length of the workpiece and here, in common parlance, referred to as directionally solidified) or a monocrystalline structure, ie the whole Workpiece consists of a single crystal. In these processes, it is necessary to avoid the transition to globulitic (polycrystalline) solidification, since non-directional growth necessarily produces transverse and longitudinal grain boundaries which negate the good properties of the directionally solidified or monocrystalline component.

The term generally refers to directionally solidified microstructures, which means both single crystals that have no grain boundaries or at most small angle grain boundaries, and stem crystal structures that have probably longitudinal grain boundaries but no transverse grain boundaries. These second-mentioned crystalline structures are also known as directionally solidified structures.

Such methods are known from U.S. Patent 6,024,792 and the EP 0 892 090 A1 known.

Likewise, the blades 120, 130 may have protection layers 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) is formed on the MCrAlX layer (as an intermediate layer or as the outermost layer).

On the MCrAlX may still be present a thermal barrier coating, which is preferably the outermost layer, and consists for example of ZrO ₂ , Y ₂ O ₃ -ZrO ₂ , ie it is not, partially or completely stabilized by yttria and / or calcium oxide and / or magnesium oxide.

The thermal barrier coating covers the entire MCrAlX layer. By suitable coating methods, e.g. Electron beam evaporation (EB-PVD) produces stalk-shaped grains in the thermal barrier coating.

Other coating methods are conceivable, e.g. atmospheric plasma spraying (APS), LPPS, VPS or CVD. The thermal barrier coating may have porous, micro- or macro-cracked grains for better thermal shock resistance. The thermal barrier coating is therefore preferably more porous than the MCrAlX layer.

The blade 120, 130 may be hollow or solid.

If the blade 120, 130 is to be cooled, it is hollow and may still film cooling holes 418 (indicated by dashed lines) on.

The FIG. 5 shows a combustion chamber 110 of the gas turbine 100. The combustion chamber 110 is configured, for example, as a so-called annular combustion chamber, in which a plurality of circumferentially arranged around a rotation axis 102 around burners 107 open into a common combustion chamber space 154, the flames 156 produce. For this purpose, the combustion chamber 110 is configured in its entirety as an annular structure, which is positioned around the axis of rotation 102 around.

To achieve a comparatively high efficiency, the combustion chamber 110 is designed for a comparatively high temperature of the working medium M of about 1000 ° C to 1600 ° C. In order to enable a comparatively long service life even with these, for the materials unfavorable operating parameters, the combustion chamber wall 153 is provided on its side facing the working medium M side with an inner lining formed from heat shield elements 155.

Due to the high temperatures inside the combustion chamber 110 may also be provided for the heat shield elements 155 and for their holding elements, a cooling system. The heat shield elements 155 are then, for example, hollow and possibly still have cooling holes (not shown) which open into the combustion chamber space 154.

Each heat shield element 155 made of an alloy is equipped on the working medium side with a particularly heat-resistant protective layer (MCrAlX layer and / or ceramic coating) or is made of high-temperature-resistant material (solid ceramic blocks).

These protective layers 7 may be similar to the turbine blades.

On the MCrAlX, for example, a ceramic thermal barrier coating may be present and consists for example of ZrO ₂ , Y ₂ O ₃ -ZrO ₂ , ie it is not, partially or completely stabilized by yttria and / or calcium oxide and / or magnesium oxide.

By suitable coating methods, e.g. Electron beam evaporation (EB-PVD) produces stalk-shaped grains in the thermal barrier coating.

Other coating methods are conceivable, e.g. atmospheric plasma spraying (APS), LPPS, VPS or CVD. The thermal barrier coating may have porous, micro- or macro-cracked grains for better thermal shock resistance.

Refurbishment means that turbine blades 120, 130, heat shield elements 155 may need to be deprotected (e.g., by sandblasting) after use. This is followed by removal of the corrosion and / or oxidation layers or products. Optionally, cracks in the turbine blade 120, 130 or the heat shield element 155 are also repaired. This is followed by a re-coating of the turbine blades 120, 130, heat shield elements 155 and a renewed use of the turbine blades 120, 130 or the heat shield elements 155.

Claims (11)

alloy
contains the following elements
(Data in wt%): 18% - 22% cobalt (Co), in particular 19% - 21%, 22% - 26% chromium (Cr), in particular 23% - 25%, 6% - 8% aluminum (Al), especially 6.5% - 7.5% 0.5% -0.7% of at least one element from the group comprising scandium and the elements of the rare earths, especially yttrium (Y),
Nickel,
especially the rest of nickel. Alloy according to claim 1,
containing 24wt% chromium. Alloy according to claim 1 or 2,
containing 20% cobalt (Co). Alloy according to one or more of Claims 1 to 3, containing 0.6% by weight of yttrium, Alloy according to one or more of claims 1 to 4, containing 7wt% aluminum. Alloy according to one or more of the preceding claims,
not containing rhenium. Alloy according to one or more of the preceding claims,
not containing silicon (Si). 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). Alloy according to one or more of the preceding claims,
consisting of cobalt (Co), chromium (Cr), aluminum (Al), yttrium (Y), nickel (Ni). Protective layer for protection of a component (1) against corrosion and / or oxidation,
especially at high temperatures,
which has the composition of the alloy according to one or more of claims 1 to 9 and
which is present in particular as a single layer. 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, for protection against corrosion and oxidation at high temperatures, has a protective layer (7) according to claim 10, in particular only a metallic protective layer (7), and
in which, in particular on the protective layer (7), a ceramic thermal barrier coating (10) is applied.

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