EP1834004A1

EP1834004A1 - Alloy, protective layer for protecting a component against corrosion and oxidation at high temperatures and component

Info

Publication number: EP1834004A1
Application number: EP06830208A
Authority: EP
Inventors: Werner Stamm
Original assignee: Siemens AG
Current assignee: Siemens AG
Priority date: 2005-12-02
Filing date: 2006-11-30
Publication date: 2007-09-19
Also published as: US20090155120A1; WO2007063091A1; KR20070099675A; EP1793008A1; 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% aluminium, 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

Alloy, protective layer for protecting a component against corrosion and / or oxidation at high temperatures and

component

The invention relates to an alloy according to claim 1, a protective layer for protecting a component against corrosion and / or oxidation at high temperatures according to claim 14 and a component according to claim 15.

The invention relates in particular to a protective layer for ^¬ a component made of a superalloy based on nickel or cobalt base.

Protective coatings for metallic components to increase their corrosion- ^¬ onsbeständigkeit and / or oxidation resistance are known in the art in great number. Most of these protective layers are known under the collective name MCrAlX, where M stands for at least one of the elements from the group consisting of iron, cobalt and nickel and wei ^¬ tere essential components chromium, aluminum and X = Ytt ^¬ rium, the latter also wholly or partially can be replaced by an equivalent element which ^¬ sem from the group comprising scandium and the rare earth elements are.

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

US Pat. No. 6,280,857 B1 discloses a protective layer which discloses the elements cobalt, chromium and aluminum based on nickel, the optional addition of rhenium as well as compelling admixtures of yttrium and silicon.

EP 1 439 245 A1 discloses a rhenium-containing cobalt-based layer. The effort to increase the inlet temperatures both in stationary gas turbines and in aircraft engines has in the field of gas turbines of great importance, since the inlet temperatures are important determinants of the achievable with gas turbines thermodynamic efficiencies. Especially through the use ent ^¬ Wrapped alloys as base materials for thermally highly onerous to components such as vanes and blades, stakes single crystal superalloy particularly through the use, inlet temperatures are possible significantly higher than 1000 ⁰ C. Meanwhile, the prior art allows inlet temperatures of 95O ⁰ C and more in stationary gas turbines and HOO ⁰ C and more in gas turbines of aircraft engines.

Examples of the construction of a turbine blade with a einkri ^¬ stable substrate, which in turn can be complex, are apparent from WO 91/01433 Al.

While the physical capacity of the now escape ^¬ oped basic materials is largely unproblematic for highly stressed components with a view to possible further increases in the inlet temperatures to resort to protective layers need to achieve sufficient resistance to oxidation and corrosion. 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 ⁰ 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. Here, the typical problem is that an increase in the An ^¬ parts of elements such as aluminum and chromium, the resistance of a protective layer against oxidation and corrosion can lead to a deterioration of the ductility of the protective layer, so that mechanical ^deformation , in particular the formation of cracks, is to be expected in the case of a mechanical load normally occurring in a gas turbine.

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 14.

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

The object is achieved by a component according to claim 15, 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.

In the subclaims further advantageous measures are listed.

The measures listed in the dependent claims can be combined in any advantageous manner with each other.

The invention is u. a. the realization that the

Protective layer in the layer and in the transition region between protective layer and base material brittle precipitates shows. These are used with time and temperature in use. Strengthens forming brittle phases lead in operation 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 precipitates increases as a result of the interaction with carbon which can diffuse from the base material into the layer or diffuses through the surface into the layer during a heat treatment in the furnace. By oxidation of the brittle precipitations, the driving force for cracking is further enhanced.

It is important the influence of nickel, which thermi ^¬ rule and determines mechanical properties.

The invention will be explained in more detail below.

Show it

1 shows a layer system with a protective layer, FIG. 2 shows compositions of superalloys, FIG. 3 shows a gas turbine,

Figure 5 is a perspective view of a turbine blade and Figure 4 is a perspective view of a combustion chamber.

According to the invention, a protective layer 7 (FIG. 1) for protecting a component against corrosion and oxidation at a high temperature comprises the following elements (indication of the parts in wt%): 2 277 %% bbiiss 31% nickel

23% to 29% chromium

7% to 11% aluminum

0.5% to 0.7% yttrium and / or at least one metal from the group comprising scandium and the elements of the rare earths, optionally 0.6% to 0.8% silicon and / or optionally 0.5% to 0, 7% zirconium, balance cobalt (CoNiCrAlY). Preferably, either only silicon or zirconium is added.

Particular embodiments are:

1) Co - 30Ni - 28Cr - 8Al - 0.6Y

2) Co - 30Ni - 28Cr - 8Al - 0.6Y - 0.7Si

3) Co - 28Ni - 24Cr - 10A1 - 0.6Y - 0.6Zr.

It should be noted that the proportions of the individual elements are specially tuned with regard to their effects. The selection of 27wt% to 31wt% nickel improves überra ^¬ dominant clearly and proportional the thermal and mechanical properties of the protective layer 7. In conjunction with the reduction of the brittle phases, which have a negative effect particularly under higher mechanical properties are represented by the reduction of the mechanical stresses improved by the selected nickel content, the mechanical properties.

The protective layer, with good corrosion resistance, has 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 with a further increase in the inlet temperature. During operation, embrittlement hardly occurs.

An equally important role is played by the trace elements in the powder to be sprayed, which form precipitates and thus embrittle them.

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 Haftvermitt ^¬ Lersch layer to a superalloy.

In this protective layer 7 may further layers insbeson ^¬ particular ceramic thermal barrier layers are deposited 10th In this device, the protective layer 7 is advantageous ^¬ as applied to a substrate 4 made of a superalloy tion on nickel or cobalt base. In particular the following composition is suitable as substrate (data in wt%):

0.1% to 0.15% carbon

18% to 22% chrome

1188 %% bbiiss 1199 %% cobalt

0% to 2% tungsten

0% to 4% molybdenum

0% to 1.5% tantalum

0% to 1% niobium

11 %% bbiiss 33 %% aluminum

2% to 4% titanium

0% 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 names GTD222, IN939, IN6203 and Udimet known 500 be ^¬.

Further alternatives for the substrate of the component are listed in FIG.

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 a component against corrosion and oxidation, while the component is acted upon by a flue gas at a material temperature of about 95O ⁰ C, in airplanes also by about HOO ⁰ 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 another component that is exposed to hot gas before or in the turbine of the gas 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).

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. In particular with turbine components, such as e.g. Turbine barrel 120 (FIG. 5) or vanes 130 (FIGS. 3, 5), combustor liners 155 (FIG. 4), as well as other housing parts of a steam or gas turbine 100 (FIG. 3), the substrate 4 is made of nickel alloy. , cobalt or iron based superalloy.

Preferably, cobalt-based or nickel-based superalloys are used. On the substrate 4, the protective layer 7 according to the invention is present.

Preferably, this protective layer is applied by LPPS (low pres sure ^¬ plasma spraying). 7

This can be used as outer layer (not shown) or ^intermediate layer (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 manufactured components and remanufactured components are introduced from the refurbishment be ^¬. Reprocessing (Refurbishment) means that components 1 are separated after their use, if appropriate, of layers (heat insulating layer) and corrosion and oxidation products are removed, for example by a Säurebehand ^¬ averaging (acid stripping). If necessary, cracks must still be repaired. Thereafter, such a component can be coated again because the substrate 4 is very expensive. FIG. 3 shows by way of example a gas turbine 100 in a longitudinal partial section.

The gas turbine 100 has an axis by a rotational ^¬ 102 rotatably mounted rotor 103 having a shaft 102, which is also referred to as the turbine rotor.

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 100 104 air 135 is sucked by the compressor 105 through the intake housing and ver ^¬ seals. The be at the turbine end of the compressor 105 ^¬ compressed air provided 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. The working medium 113 expands on the rotor blades 120 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 a 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 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 are part of the disclosure regarding the chemical composition of the alloys.

The guide vane 130 has an inner housing 138 of the turbine 108 facing guide vane root (not provide Darge ^¬ here) and a side opposite the guide-blade root vane root. The vane head faces the rotor 103 and fixed to a mounting ring 140 of the stator 143.

5 shows a perspective view of a rotor blade 120 or guide vane show ^¬ 130 of a turbomachine, which extends along a longitudinal axis of the 121st

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 to each other, a securing region 400, an adjoining blade or vane platform 403 and a blade 406 and a blade tip 415. As a guide vane 130, the vane 130 may have tip at its ^vane 415 a further platform (not Darge ^¬ asserted).

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, for example, as a hammerhead staltet out ^¬. Other designs as Christmas tree or Schwalbenschwanzfuß are possible. The blade 120, 130 has a medium which flows past felblatt 406 at the spectacle ^¬ a leading edge 409 and a ^trailing edge on the 412th

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

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; These documents are part of the disclosure regarding the chemical composition of the alloy. The blade 120, 130 can hereby be manufactured 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 production of such monocrystalline workpieces, for example, by directed solidification from the melt. These are casting processes in which the liquid metallic alloy to monocrystalline structure, ie 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 be ^¬ is made of a single crystal. In these methods, you have to transition to globular (polycrystalline) solidification avoided, since non-directional growth inevitably forms transverse and longitudinal grain boundaries ^¬ which the good properties of the directionally solidified or single-crystal component nullify. If the term generally refers to directionally solidified structures, it means both single crystals that have no grain boundaries or at most small-angle grain boundaries, and stem crystal structures that have grain boundaries that are probably longitudinal but no transverse grain boundaries. In these second-mentioned crystalline

Structures are also known as directionally rigidified structures

(directionally solidified structures).

Such methods are known from US Pat. No. 6,024,792 and EP

0 892 090 Al known; these writings are part of the disclosure with regard to the solidification process.

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 fully stabilized by yttria and / or calcium oxide and / or magnesium oxide. The thermal barrier coating covers the entire MCrAlX layer. Suitable coating processes, such as electron beam evaporation (EB-PVD), produce stalk-shaped grains in the thermal barrier coating. Other coating methods are conceivable, for example atmospheric plasma spraying (APS), LPPS, VPS or CVD. The heat insulating layer can ^¬ porous, micro- or macro-cracked ^compatible grains have 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 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 designed, for example, as a so-called annular combustion chamber, in which a multiplicity of burners 107 arranged in the circumferential direction about a rotation axis 102 open into a common combustion chamber space 154, the flames 156 generate. 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. A relatively long service life even with these handy, unfavorable for the materials to operational parameters made ^¬, the combustion chamber wall 153 is provided on its side facing the ^working medium M facing side with a formed from heat shield elements 155. liner.

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 hollow, for example, 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 fluid 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 120, 130, so for example MCrAlX means: M is at least one element of the group iron (Fe), cobalt (Co), nickel (Ni), X is an active element and stands for yttrium (Y) and / or silicon and / or at least one element of the rare earths, or hafnium (Hf). Such alloys are known from EP 0 486 489 B1, EP 0 786 017 Bl, EP 0 412 397 B1 or EP 1 306 454 A1.

On the MCrAlX, for example, a ceramic thermal insulation layer may be present and consists for example of ZrO ₂ , Y ₂ O ₃ -ZrO ₂ , ie it is not, partially or fully ^¬ dig stabilized by yttrium 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, for example atmospheric plasma spraying (APS), LPPS, VPS or CVD. The heat insulation layer ^¬ can porous, micro- or macro-cracked pERSonal for better thermal shock resistance.

Reprocessing (Refurbishment) means that turbines ^¬ blades 120, 130, heat shield elements 155 after they set ^¬ A may need to be stripped of protective layers (for example by sandblasting). 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. After that 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

claims

1. Alloy containing the following elements

(Data 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 from the group comprising scandium and the elements of

Rare earth optional 0.6% to 0.8% silicon optional 0.5% to 0.7% zirconium residual cobalt.

2. An alloy according to claim 1, containing 27wt% to 29wt% nickel, in particular 28wt% nickel.

3. An alloy according to claim 1, containing 29wt% to 31wt% nickel, in particular 30wt% nickel.

4. The alloy of claim 1, 2 or 3, containing 23wt% to 25wt% chromium, in particular 24 wt% chromium.

5. The alloy of claim 1, 2 or 3, containing 27wt% to 29wt% chromium, in particular 28wt% chromium.

6. The alloy of claim 1, 2, 3, 4 or 5, containing 7wt% to 9wt% aluminum, in particular 8wt% aluminum.

7. The alloy of claim 1, 2, 3, 4 or 5, containing 9wt% to llwt% aluminum, in particular 10wt% aluminum.

8. Alloy according to one or more of claims 1 to 7, containing 0.6% yttrium.

9. Alloy according to one or more of claims 1 to containing silicon, in particular 0.7% silicon.

10. Alloy according to one or more of claims 1 to 9, containing zirconium, in particular 0.6% zirconium.

11. An alloy according to one or more of claims 1 to 10, consisting of cobalt, nickel, aluminum, chromium, yttrium and optionally zirconium and / or silicon.

12. An alloy according to claim 1, 9 or 11, containing silicon, not containing zirconium.

13. An alloy according to claim 1, 10 or 11, containing zirconium, not containing silicon.

14. protective layer (7) for protecting a component (1, 120,

130, 155) 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 13.

15. component, in particular a component (120, 130, 155) of a gas turbine

(100), which protects against corrosion and oxidation at high

Temperatures a protective layer (7) according to claim 14 has.

16. Component according to claim 15, wherein on the protective layer (7) a thermal barrier coating (10) is applied.

17. Component according to claim 15 or 16, wherein a substrate (4) of the component (120, 130, 155) is nickel-based.

18. Component according to claim 15 or 16, wherein a substrate (4) of the component (120, 130, 155) is cobalt-based.