US20090155120A1

US20090155120A1 - Alloy, Protective Layer for Protecting a Component Against Corrosion and/or Oxidation at High Temperatures, and Component

Info

Publication number: US20090155120A1
Application number: US12/085,701
Authority: US
Inventors: Werner Stamm
Original assignee: Siemens AG
Current assignee: Siemens AG
Priority date: 2005-12-02
Filing date: 2006-11-30
Publication date: 2009-06-18
Also published as: EP1793008A1; EP1834004A1; WO2007063091A1; KR20070099675A; CN101133173A

Abstract

Known protective layers with a high Cr content and additionally a silicon form brittle phases, which become even more brittle under the influence of carbon during use. The protective layer according to the invention has the composition 27% to 31% nickel, 23% to 29% chromium, 7% to 11% aluminum, 0.5% to 0.7% yttrium and/or at least one equivalent metal from the group comprising scandium and rare earth elements, optionally 0.6% to 0.8% silicon, optionally 0.5% to 0.7% zirconium and the remainder cobalt.

Description

CROSS REFERENCE TO RELATED APPLICATIONS

This application is the US National Stage of International Application No. PCT/EP2006/069104, filed Nov. 30, 2006 and claims the benefit thereof. The International Application claims the benefits of European application No. 05026378.9 filed Dec. 2, 2005, both of the applications are incorporated by reference herein in their entirety.

FIELD OF INVENTION

The invention relates to an alloy as described in the claims, a protective layer for protecting a component against corrosion and/or oxidation at high temperatures and a component as described in the claims.
The invention relates in particular to a protective layer for a component that consists of a nickel-base or cobalt-base superalloy.

BACKGROUND OF THE INVENTION

Numerous protective layers for metallic components that are supposed to increase the corrosion resistance and/or oxidation resistance of said components are known from the prior art. Most of these protective layers are known under the collective name MCrAlX, where M stands for at least one of the elements selected from the group consisting of iron, cobalt and nickel and further essential constituents are chromium, aluminum and X=Yttrium, wherein the latter may also be partially or completely replaced by an equivalent element selected from the group consisting of scandium and the rare earth elements.
Typical coatings of this type are known from U.S. Pat. Nos. 4,005,989 and 4,034,142.
U.S. Pat. No. 6,280,857 B1 discloses a protective layer which contains the elements cobalt, chromium and aluminum, based on nickel, with the optional addition of rhenium and obligatory admixtures of yttrium and silicon.
EP 1 439 245 A1 discloses a cobalt-based rhenium-containing layer.
The objective of increasing the inlet temperatures of both stationery gas turbines and aircraft engines is of considerable significance in the specialist field of gas turbines, since the inlet temperatures are important variables determining the thermodynamic efficiencies which can be achieved by gas turbines. The use of specially developed alloys as base materials for components which are to be exposed to high thermal stresses, such as guide vanes and rotor blades, and in particular the use of single-crystal superalloys, allows the use of inlet temperatures of well over 1000° C. Nowadays, the prior art permits inlet temperatures of 950° C. and above in the case of stationary gas turbines and 1100° C. and above in the case of gas turbines for aircraft engines.
Examples of the structure of a turbine blade or vane having a single-crystal substrate, which for its part may be of complex structure, are revealed by WO 91/01433 A1.
Whereas the physical load-bearing capacity of the base materials which have by now been developed for the highly stressed components does not present any major problems with a view to possible further increases in the inlet temperatures, protective layers have to be employed to achieve sufficient resistance to oxidation and corrosion. In addition to the sufficient chemical stability of a protective layer under the attacks expected from flue gases at temperatures of the order of magnitude of 1000° C., a protective layer also has to have sufficiently good mechanical properties, not least with a view to the mechanical interaction between the protective layer and the base material. In particular, the protective layer must be sufficiently ductile to enable any deformation of the base material to be followed and not to crack, since points of attack for oxidation and corrosion would be created in this way. This typically gives rise to the problem that an increase in the levels of elements such as aluminum and chromium, which can increase the resistance of a protective layer to oxidation and corrosion, leads to a deterioration in the ductility of the protective layer, which means that mechanical failure, in particular the formation of cracks, is likely under mechanical loading which usually occurs in a gas turbine.

SUMMARY OF INVENTION

Accordingly, the invention is based on the object of providing an alloy and a protective layer which has a good high-temperature stability with regard to corrosion and oxidation, good long-term stability and, moreover, is particularly well matched to mechanical stresses which are expected at a high temperature in particular in a gas turbine.
The object is achieved by the alloy as claimed in the claims and the protective layer as claimed in the claims.
A further object of the invention is to provide a component which offers increased protection against corrosion and oxidation.
This object is achieved by the component as claimed in the claims, in particular a component of a gas turbine or steam turbine, which for protection against corrosion and oxidation at high temperatures, has a protective layer of the type described above.
The subclaims list further advantageous measures.
The measures listed in the subclaims can be combined with one another as desired in advantageous ways.
The invention is based on the discovery, inter alia, that the desired protective layer has brittle precipitates in the layer and also in the transition region between the protective layer and the base material. These brittle phases, the formation of which increases over time and with use temperature, in operation lead to highly pronounced longitudinal cracks in the layer and in the layer/base material interface, with subsequent layer detachment. The interaction with carbon, which can diffuse out of the base material into the layer or diffuses into the layer through the surface during a heat treatment in the furnace, additionally increases the brittleness of the precipitates. The susceptibility to cracking is boosted still further by oxidation of the brittle precipitates.
In this context, the influence of nickel, which determines thermal and mechanical properties, is also important.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention is explained in more detail below. In the drawing:

FIG. 1 shows a layer system having a protective layer,

FIG. 2 shows compositions of superalloys,

FIG. 3 shows a gas turbine,

FIG. 5 shows a perspective view of a turbine blade or vane, and

FIG. 4 shows a perspective view of a combustion chamber.

DETAILED DESCRIPTION OF INVENTION

According to the invention, a protective layer 7 (FIG. 1) for protecting a component against corrosion and oxidation at high temperature comprises the following elements (details of amounts in wt %):
27% to 31% nickel
23% to 29% chromium
7% to 11% aluminum
0.5% to 0.7% yttrium and/or at least one metal selected from the group consisting of scandium and the rare earth elements, optionally 0.6% to 0.8% silicon, and/or optionally 0.5% to 0.7% zirconium, remainder cobalt (CoNiCrAlY).
It is preferable for either only silicon or zirconium to be added.
Particular exemplary embodiments are:
1) Co-30Ni-28Cr-8Al-0.6Y
2) Co-30Ni-28Cr-8Al-0.6Y-0.7Si
3) Co-28Ni-24Cr-10Al-0.6Y-0.6Zr.
It should be noted that the levels of the individual elements are specifically adapted with a view to their actions. Surprisingly, the selection of 27 wt % to 31 wt % nickel significantly and disproportionately improves the thermal and mechanical properties of the protective layer 7.
In conjunction with the reduction in brittle phases, which have negative effects in particular with relatively elevated mechanical properties, the reduction in the mechanical stresses resulting from the selected nickel content improves the mechanical properties.
The protective layer, with a good resistance to corrosion, has a particularly good resistance to oxidation and is furthermore distinguished by especially good ductility properties, making it particularly well qualified for use in a gas turbine with a further increase in the inlet temperature. Scarcely any embrittlement occurs during operation.
The trace elements in the powder to be sprayed, which form precipitates and therefore constitute sources of embrittlement, also play an important role.
The powders are applied, for example, by plasma spraying (APS, LPPS, VPS, . . . ). Other processes are also conceivable (PVP, CVD, cold spraying).
The protective layer 7 described also acts as a bonding layer to a superalloy.
Further layers, in particular ceramic thermal barrier coatings 10, can be applied to this protective layer 7.
In this component, the protective layer 7 is advantageously applied to a substrate 4 made from a nickel-base or cobalt-base superalloy.
A suitable substrate has in particular the following composition (details in wt %):


0.1% to 0.15%	Carbon
18% to 22%	Chromium
18% to 19%	Cobalt
0% to 2%	Tungsten
0% to 4%	Molybdenum
0% to 1.5%	Tantalum
0% to 1%	Niobium
1% to 3%	Aluminum
2% to 4%	Titanium
0% to 0.75%	Hafnium

optionally small quantities of boron and/or zirconium, remainder nickel.
Compositions of this type are known as casting alloys under the names GTD222, IN939, IN6203 and Udimet 500.
Further alternatives for the substrate of the component are listed in FIG. 2.
The thickness of the protective layer 7 on the component 1 is preferably between approximately 100 μm and 300 μm.
The protective layer 7 is particularly suitable for protecting a component against corrosion and oxidation when the component is exposed to a flue gas at a material temperature of around 950° C., and in the case of aircraft turbines even around 1100° C.
The protective layer 7 according to the invention is therefore particularly well qualified for protecting a component of a gas turbine 100, in particular a guide vane 120, rotor blade 130 or other component, which is exposed to hot gas upstream of or in the turbine of the gas turbine.
The protective layer 7 can be used as an overlay (the protective layer is the outer layer) or as a bond coat (the protective layer is an interlayer).
FIG. 1 shows a layer system 1 as a component.
The layer system 1 comprises a substrate 4.
The substrate 4 may be metallic and/or ceramic. In particular in the case of turbine components, such as for example turbine rotor blades 120 (FIG. 5) or turbine guide vanes 130 (FIGS. 3, 5), combustion chamber linings 155 (FIG. 4) and other housing parts of a steam or gas turbine 100 (FIG. 3), the substrate 4 consists of a nickel-base, cobalt-base or iron-base superalloy.
It is preferable to use cobalt-base or nickel-base superalloys.
The protective layer 7 according to the invention is present on the substrate 4.
It is preferable for this protective layer 7 to be applied by LPPS (low pressure plasma spraying).
It can be used as the outer layer (not shown) or as the interlayer (FIG. 1).
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 produced components and refurbished components.
Refurbishment means that after they have been used, layers (thermal barrier coating) may have to be detached from components 1 and corrosion and oxidation products removed, for example by an acid treatment (acid stripping). If appropriate, cracks also have to be repaired. This can be followed by recoating of a component of this type, since the substrate 4 is very expensive.
FIG. 3 shows by way of example a partial longitudinal section through a gas turbine 100.
In its interior, the gas turbine 100 has a rotor 103 which is mounted such that it can rotate about an axis of rotation 102, has a shaft 102, and is also referred to as the turbine rotor.
An intake casing 104, a compressor 105, a for example toric combustion chamber 110, in particular an annular combustion chamber, with a plurality of coaxially arranged burners 107, a turbine 108 and the exhaust gas casing 109 follow one another along the rotor 103.
The annular combustion chamber 110 is in communication with a for example annular hot gas duct 111. There, by way of example, four successive turbine stages 112 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, a guide vane row 115 is followed in the hot gas duct 111 by a row 125 formed from rotor blades 120.
The guide vanes 130 are secured to an inner casing 138 of a stator 143, whereas the rotor blades 120 belonging to a row 125 are arranged on the rotor 103, for example by means of a turbine disk 133.
A generator (not shown) is coupled to the rotor 103.
While the gas turbine 100 is operating, air 135 is drawn in through the intake casing 104 and compressed by the compressor 105. The compressed air provided at the turbine end of the compressor 105 is passed to the burners 107, where it is mixed with a fuel. The mixture is then burnt in the combustion chamber 110, forming 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. The working medium 113 is expanded at the rotor blades 120, transferring its momentum, so that the rotor blades 120 drive the rotor 103 and the latter in turn drives the generator coupled to it.
While the gas turbine 100 is operating, the components which are exposed to the hot working medium 113 are subject to thermal stresses. The guide vanes 130 and rotor blades 120 of the first turbine stage 112, as seen in the direction of flow of the working medium 113, together with the heat shield elements which line the annular combustion chamber 110, are subject to the highest thermal stresses.
To be able to withstand the temperatures which prevail there, they can be cooled by means of a coolant.
Substrates of the components may likewise have a directional structure, i.e. they are in single-crystal form (SX structure) or have only longitudinally oriented grains (DS structure).
By way of example, iron-base, nickel-base or cobalt-base superalloys are used as material for the components, in particular for the turbine blade or vane 120, 130 and components of the combustion chamber 110.
Superalloys of this type are known for example from EP 1 204 776 B1, EP 1 306 454, EP 1 319 729 A1, WO 99/67435 or WO60/44949; these documents form part of the disclosure with regard to the chemical composition of the alloys.
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 head at the opposite end from the guide vane root. The guide vane head faces the rotor 103 and is fixed to a securing ring 140 of the stator 143.
FIG. 5 shows a perspective view of a rotor 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 of a power plant for generating electricity, a steam turbine or a compressor.
The blade or vane 120, 130 has, in succession along the longitudinal axis 121, a securing region 400, an adjoining blade or vane platform 403, a main blade or vane part 406 and a blade or vane tip 415.
As a guide vane 130, the vane 130 may have a further platform (not shown) at its vane tip 415.
A blade or vane root 183, which is used to secure the rotor blades 120, 130 to a shaft or a disk (not shown), is formed in the securing region 400.
The blade or vane root 183 is designed, for example, in hammerhead form. Other configurations, such as a fir-tree or dovetail root, are possible.
The blade or vane 120, 130 has a leading edge 409 and a trailing edge 412 for a medium which flows past the main blade or vane part 406.
In the case of conventional blades or vanes 120, 130, by way of example solid metallic materials, in particular superalloys, are used in all regions 400, 403, 406 of the blade or vane 120, 130.
Superalloys of this type 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; these documents form part of the disclosure with regard to the chemical composition of the alloy.
The blade or vane 120, 130 may in this case be produced 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 which, in operation, are exposed to high mechanical, thermal and/or chemical stresses.
Single-crystal workpieces of this type are produced, for example, by directional solidification from the melt. This involves casting processes in which the liquid metallic alloy solidifies to form the single-crystal structure, i.e. the single-crystal workpiece, or solidifies directionally.
In this case, dendritic crystals are oriented along the direction of heat flow and form either a columnar crystalline grain structure (i.e. grains which run over the entire length of the workpiece and are referred to here, in accordance with the language customarily used, as directionally solidified) or a single-crystal structure, i.e. the entire workpiece consists of one single crystal. In these processes, a transition to globular (polycrystalline) solidification needs to be avoided, since non-directional growth inevitably forms transverse and longitudinal grain boundaries, which negate the favorable properties of the directionally solidified or single-crystal component.
Where the text refers in general terms to directionally solidified microstructures, this is to be understood as meaning both single crystals, which do not have any grain boundaries or at most have small-angle grain boundaries, and columnar crystal structures, which do have grain boundaries running in the longitudinal direction but do not have any transverse grain boundaries. This second form of crystalline structures is also described as directionally solidified microstructures. (directionally solidified structures).
Processes of this type are known from U.S. Pat. No. 6,024,792 and EP 0 892 090 A1; these documents form part of the disclosure with regard to the solidification process.
The blades or vanes 120, 130 may likewise have protective layers 7 according to the invention protecting against corrosion or oxidation. The density is preferably 95% of the theoretical density. A protective aluminum oxide layer (TGO=thermally grown oxide layer) is formed on the MCrAlX layer (as an interlayer or as the outermost layer).
It is also possible for a thermal barrier coating, which is preferably the outermost layer and consists for example of ZrO₂, Y₂O₃—ZrO₂, i.e. unstabilized, partially stabilized or fully stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide, to be present on the MCrAlX.
The thermal barrier coating covers the entire MCrAlX layer.
Columnar grains are produced in the thermal barrier coating by means of suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD).
Other coating processes are conceivable, for example atmospheric plasma spraying (APS), LPPS, VPS or CVD. The thermal barrier coating may have grains that are porous and/or include micro-cracks or macro-cracks in order to improve the resistance to thermal shocks. Therefore, the thermal barrier coating is preferably more porous than the MCrAlX layer.
The blade or vane 120, 130 may be hollow or solid in form. If the blade or vane 120, 130 is to be cooled, it is hollow and may also have film-cooling holes 418 (indicated by dashed lines).
FIG. 4 shows a combustion chamber 110 of the gas turbine 100. The combustion chamber 110 is configured, for example, as what is known as an annular combustion chamber, in which a multiplicity of burners 107, which generate flames 156 and are arranged circumferentially around an axis of rotation 102, open out into a common combustion chamber space 154. For this purpose, the combustion chamber 110 overall is of annular configuration positioned around the axis of rotation 102.
To achieve a relatively high efficiency, the combustion chamber 110 is designed for a relatively high temperature of the working medium M of approximately 1000° C. to 1600° C. To allow a relatively long service life even with these operating parameters, which are unfavorable for the materials, the combustion chamber wall 153 is provided, on its side which faces the working medium M, with an inner lining formed from heat shield elements 155.
A cooling system may also be provided for the heat shield elements 155 and/or their holding elements, on account of the high temperatures in the interior of the combustion chamber 110. The heat shield elements 155 are then for example hollow and may also have cooling holes (not shown) which open out into the combustion chamber space 154.
On the working medium side, each heat shield element 155 made from an alloy is equipped with a particularly heat-resistant protective layer (MCrAlX layer and/or ceramic coating) or is made from material that is able to withstand high temperatures (solid ceramic bricks).
These protective layers 7 may be similar to those used for the turbine blades or vanes 120, 130, i.e. for example MCrAlX: M is at least one element selected from the group consisting of iron (Fe), cobalt (Co), nickel (Ni), X is an active element and stands for yttrium (Y) and/or silicon and/or at least one of the rare earth elements, or hafnium (Hf). Alloys of this type are known from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1.
A for example ceramic thermal barrier coating, consisting for example of ZrO₂, Y₂O₃—ZrO₂, i.e. unstabilized, partially stabilized or fully stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide, may also be present on the MCrAlX.
Columnar grains are produced in the thermal barrier coating by suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD).
Other coating processes are conceivable, for example atmospheric plasma spraying (APS), LPPS, VPS or CVD. The thermal barrier coating may have grains that are porous and/or include micro-cracks or macro-cracks in order to improve the resistance to thermal shocks.
Refurbishment means that after they have been used, protective layers may have to be removed from turbine blades or vanes 120, 130, heat shield elements 155 (e.g. by sand-blasting). Then, the corrosion and/or oxidation layers and products are removed.
If appropriate, cracks in the turbine blade or vane 120, 130 or the heat shield element 155 are also repaired. This is followed by recoating of the turbine blades or vanes 120, 130, heat shield elements 155, after which the turbine blades or vanes 120, 130 or the heat shield elements 155 can be reused.

Claims

1.-18. (canceled)

19. An alloy, comprising (in wt %):

27% to 31% nickel;

23% to 29% chromium;

7% to 11% aluminum;

0.5% to 0.7% yttrium or at least one metal selected from the group consisting of scandium and the rare earth elements;

optionally 0.6% to 0.8% silicon;

optionally 0.5% to 0.7% zirconium; and

remainder cobalt.

20. The alloy as claimed in claim 19, wherein the alloy comprises 28 wt % nickel.

21. The alloy as claimed in claim 19, wherein the alloy comprises 30 wt % nickel.

22. The alloy as claimed in claim 21, wherein the alloy comprises 24 wt % chromium.

23. The alloy as claimed in claim 21, wherein the alloy comprises 28 wt % chromium.

24. The alloy as claimed in claim 23, wherein the alloy comprises 8 wt % aluminum.

25. The alloy as claimed in claim 23, wherein the alloy comprises 10 wt % aluminum.

26. The alloy as claimed in claim 25, wherein the alloy comprises 0.6 wt % yttrium.

27. The alloy as claimed in claim 26, wherein the alloy comprises 0.7% silicon.

28. The alloy as claimed in claim 27, wherein the alloy comprises 0.6% zirconium.

29. The alloy as claimed in claim 28, wherein the alloy comprises silicon and not zirconium.

30. The alloy as claimed in claim 28, wherein the alloy comprises zirconium and not silicon.

31. A protective layer for protecting a component against corrosion and oxidation at high temperatures, comprising (in wt %):

27% to 31% nickel;

23% to 29% chromium;

7% to 11% aluminum;

optionally 0.6% to 0.8% silicon;

optionally 0.5% to 0.7% zirconium; and

remainder cobalt.

32. A component of a gas turbine having a protective layer for protection against corrosion and oxidation at high temperatures, comprising (in wt %):

27% to 31% nickel;

23% to 29% chromium;

7% to 11% aluminum;

optionally 0.6% to 0.8% silicon;

optionally 0.5% to 0.7% zirconium; and

remainder cobalt.

33. The component as claimed in claim 32, wherein a thermal barrier coating has been applied to the protective layer.

34. The component as claimed in claim 33, wherein a substrate of the component is nickel-based.

35. The component as claimed in claim 34, wherein a substrate of the component is cobalt-based.