POROUS CERAMIC LAYER SYSTEM

Abstract

A porous ceramic layer system is provided having two layers of porous layers with a tightly controlled and matched porosity. The layer system has a substrate, a metallic bonding layer on the substrate, an inner ceramic layer having a porosity, an outer ceramic layer on the inner ceramic layer, with a two-layered NiCoCrAlX layer as the metallic bonding layer, with a lower NiCoCrAlY layer, in which the content of chromium (Cr) in the lower NiCoCrAlY layer is less than the content of chromium (Cr) in the outer NiCoCrAlY layer.

Description

CROSS REFERENCE TO RELATED APPLICATIONS

This application claims the benefit of and priority to European Application No. EP13183718 filed Sep. 10, 2013, incorporated by reference herein in its entirety.

FIELD OF INVENTION

The invention relates to a layer system having two different porous ceramic layers.

BACKGROUND OF INVENTION

Ceramic protective layers are often used for components used at high temperatures in order to protect the metallic substrate from relatively high temperatures.

In this respect, the ceramic layers have a certain porosity in order firstly to reduce the thermal conductivity and in order to set a certain ductility.

SUMMARY OF INVENTION

It is an object of the invention to optimize the thermal and the mechanical properties.

The object is achieved by a layer system as claimed.

The dependent claims list further advantageous measures which can be combined with one another, as desired.

The ceramic and metallic layers ensure good oxidation protection and good thermal insulation.

BRIEF DESCRIPTION OF THE DRAWINGS

In the drawing:

FIG. 1 shows a layer system,

FIG. 2 shows a gas turbine,

FIG. 3 shows a turbine blade or vane,

FIG. 4 shows a combustion chamber,

FIG. 5 shows a list of superalloys.

DETAILED DESCRIPTION OF INVENTION

The description and the figures show merely exemplary embodiments of the invention.

FIG. 1 schematically shows the layer system.

The layer system 1 is preferably a turbine blade or vane 120, 130 of a turbine, of a steam turbine, of a gas turbine 100 (FIG. 2), for stationary operation or for aircraft.

It is preferable for the substrate 4 to comprise a nickel-based or cobalt-based superalloy made of an alloy shown in FIG. 5. This is preferably a nickel-based superalloy.

A metallic bonding layer 7, in particular of the MCrAl or MCrAlX type (M=Ni, Co, Fe, preferably Ni, Co), is preferably present on the substrate 4.

A ceramic layer 16 is applied to the metallic bonding layer 7, wherein an oxide layer (TGO) is either produced deliberately or applied between metallic layers and ceramic layers or forms during the ceramic coating or during operation of a layer system having the metallic layer 7.

The ceramic layer 16 has at least two, in particular only two, different ceramic layers 19, 22.

The lower ceramic layer 19 has a lower porosity than the outer ceramic layer 22.

The porosity of the lower ceramic layer 19 is 8% to <15%, in particular 11% to 13% (preferably % by volume).

It is preferable that the layer thickness of the inner ceramic layer 19 is at least 10%, in particular 20%, very particularly 50% thinner than that of the outer ceramic layer 22.

The lower ceramic layer 19 has a thickness of 100±25 μm, whereas the outer ceramic layer 22 has a thickness of >100 μm.

The outer ceramic layer 22 has a porosity of >15% to 22%, in particular of 16% to 20%, and is preferably the outermost ceramic layer 22.

The material of the lower ceramic layer 19 is partially stabilized, in particular by yttrium, zirconium oxide.

This material is preferably also used for the outer ceramic layer 22, although it is also possible to use a pyrochlore material, such as gadolinium zirconate.

The selection of the porosity of the ceramic layers surprisingly led to a longer service life compared to a layer of high porosity and equal thickness.

The fact that these elements of the metallic layer 7 also interact with the base material of the substrate 4 owing to diffusion must also be taken into great consideration.

In general, it is assumed that, owing to the relatively great interdiffusion of chromium from the layer into the base material, which generally has a lower chromium content than the layer, the difference between the chromium contents in the layer and the base material should not be greater than approximately 5%. Otherwise, a more or less severe Kirkendall porosity will arise, leading to premature failure of the layer assembly with the base material. This has been confirmed by model calculations carried out appropriately. This behavior has been confirmed experimentally, as proven by the comparison of layers having a low chromium content and a high chromium content on IN 738 LC.

On the other hand, for the upper limit of the chromium content of the layer, it should be taken into account that, given low chromium contents of approximately 13% by weight chromium (Cr) in the layer, spinel formation with multiple cracking often arises at the surface, likewise leading to a shortened service life of the protective layer system. Although a very balanced composition of the protective layer already leads to good results, this does not yet constitute the optimum.

For the reasons mentioned above, a solution has been sought which combines all the advantages.

The solution proposed here presents a combination of layer compositions as a duplex layer, which, compared to layer compositions to date, has improvements in terms of the aforementioned problems.

The assertions described are shown schematically and as metallographic images in the enclosures.

What is proposed is a protective layer which, compared to the layers used to date, has better oxidation resistance and good thermomechanical properties and, on account of the substitution of rhenium, has considerable cost benefits. In addition, the interdiffusion behavior is said to be the same or better. In contrast to conventional layer compositions, the upper layer of the duplex layer has chromium contents of >20% chromium, in particular >22% chromium (Cr). This avoids spinel formation and multiple cracking in the TGO. The higher chromium (Cr) content in the topmost layer has two reasons: on the one hand, despite evaporation of chromium (Cr) during the solution annealing treatment, enough Cr remains present in the topmost layer in order to keep the activity of aluminum high, and on the other hand the chromium serves as a nucleating agent for stable alpha-aluminum oxide.

The lower layer (boundary layer to the base material) of the duplex layer has by contrast a considerably lower chromium content, preferably of approximately 11% by weight-15% by weight chromium (Cr). This prevents a Kirkendall porosity which reduces service life at the interface with the base material.

The other constituents of the layers are based on optimized proportions of nickel (Ni), cobalt (Co), aluminum (Al), rare earth elements (Y, . . . ) and the like, but no rhenium (Re).

Example: Metallic duplex protective layer 7 comprises at least: a lower NiCoCrAlY layer 10: an NiCoCrAlY protective layer having the composition (in % by weight) of Ni content: remainder cobalt (Co): 24%-26%, in particular 25%, chromium (Cr): 11%-15%, in particular 13%, aluminum (Al): 10.0%-12.0%, in particular 11.0%, yttrium (Y): 0.2%-0.6%, in particular 0.3% to 0.5%

Moderately high Co content: broadening of the beta/gamma field, avoidance of brittle phases

Moderate Cr content: low enough to avoid brittle phases (alpha-chromium or sigma phase) and to avoid Kirkendall porosity and nevertheless to preserve the protective action over long periods of time

Moderately high Al content: sufficiently high to additionally deliver Al to preserve a stable TGO. Low enough to achieve good ductility and to avoid tendency toward embrittlement

Low Y content: sufficiently high to still form sufficient Y aluminate for forming Y-containing pegs with low oxygen contamination low enough to negatively accelerate the oxide layer growth of the Al₂O₃ layer, and also an upper NiCoCrAlY layer (13): an NiCoCrAlY protective layer (13) having the composition (in % by weight) of Ni content: remainder cobalt (Co): 24%-26%, preferably 25%, chromium (Cr): 23%-25%, preferably 24%, aluminum (Al): 9%-12.0%, preferably 10.5%, yttrium (Y): 0.2%-0.6%, preferably 0.3%-0.5%

High Cr content: to avoid spinel and multiple cracking in the TGO and improve the oxide layer formation of Al₂O₃ with low oxidation rates

Moderately high Al content: the Al content is lowered slightly compared to the lower layer in order to minimize impairment of the ductility by the high Cr content.

The NiCoCrAlY layers/alloys can also comprise further elements, other or further rare earth elements or Ta, Ti, Fe, . . . , but no rhenium (Re).

No chromizing of an individual layer is carried out for the upper NiCoCrAlY layer 13, and therefore there is also no chromium gradient present, because a uniform powder is used in order to apply the layer.

Thermodynamic phase calculations and also test results for the respective individual layer have shown that good results are present in terms of oxidation, formation of the TGO and the mechanical properties.

The total layer thickness of the metallic layer 7 on the blade or vane 120, 130 should preferably be 180 μm to 300 μm.

The lower metallic layer 10 is preferably sprayed with a fine powder and the upper layer 13 consists of the powder having a high chromium content with a relatively coarse powder fraction, in order to provide not only the improved oxide layer formation but also the required high roughness of R_a=9 μm to 14 μm for optimum bonding for a ceramic layer.

This procedure also has the advantage that no new cost-increasing process step is necessary.

FIG. 1 shows a layer system consisting of a substrate 4 and the two-layered NiCoCrAlY layer 7, which is composed of two different layer compositions 10, 13.

Nickel-based or cobalt-based superalloys, in particular alloys as shown in FIG. 3, can be used as the substrate 4.

FIG. 2 shows, by way of example, a partial longitudinal section through a gas turbine 100.

In the interior, the gas turbine 100 has a rotor 103 with a shaft which is mounted such that it can rotate about an axis of rotation 102 and is also referred to as the turbine rotor.

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

The annular combustion chamber 110 is in communication with a, for example, annular hot-gas passage 111. There, four successive turbine stages 112 form the turbine 108, for example.

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

The guide vanes 130 are secured to an inner housing 138 of a stator 143, whereas the rotor blades 120 of a row 125 are fitted to 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, the compressor 105 sucks in air 135 through the intake housing 104 and compresses it. The compressed air provided at the turbine-side end of the compressor 105 is passed to the burners 107, where it is mixed with a fuel. The mix is then burnt in the combustion chamber 110, forming the working medium 113. From there, the working medium 113 flows along the hot-gas passage 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 may 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-based, nickel-based or cobalt-based 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 WO 00/44949.

The guide vane 130 has a guide vane root (not shown here), which faces the inner housing 138 of the turbine 108, and a guide vane head which is 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. 3 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 and 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.

The blade or vane 120, 130 may in this case be produced by a casting process, 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.

The blades or vanes 120, 130 may likewise have coatings protecting against corrosion or oxidation e.g. (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 rare earth element, 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.

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 intermediate layer or as the outermost layer).

The layer preferably has a composition Co-30Ni-28Cr-8Al-0.6Y-0.7Si or Co-28Ni-24Cr-10Al-0.6Y. In addition to these cobalt-based protective coatings, it is also preferable to use nickel-based protective layers, such as Ni-10Cr-12Al-0.6Y-3Re or Ni-12Co-21Cr-11Al-0.4Y-2Re or Ni-25Co-17Cr-10Al-0.4Y-1.5Re.

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 suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD).

Other coating processes are possible, e.g. atmospheric plasma spraying (APS), LPPS, VPS or CVD. The thermal barrier coating may include grains that are porous or have micro-cracks or macro-cracks, in order to improve the resistance to thermal shocks. The thermal barrier coating is therefore 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, 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.

Moreover, a cooling system may 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) opening 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 may be similar to the turbine blades or vanes, 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 rare earth element 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.

It is also possible for a for example ceramic thermal barrier coating, which 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.

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 possible, e.g. atmospheric plasma spraying (APS), LPPS, VPS or CVD. The thermal barrier coating may include grains that are porous or have 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 or 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 or 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. A layer system at least comprising:

a substrate,

a metallic bonding layer on the substrate,

an inner ceramic layer having a porosity, in % by volume, of 8% to <15% on the bonding layer,

and an outer ceramic layer, having a porosity of >15% to 22%, on the inner ceramic layer,

a two-layered NiCoCrAlX layer as the metallic bonding layer,

with a lower NiCoCrAlY layer, in which the content of chromium (Cr) in the lower NiCoCrAlY layer is at least 3% by weight less than the content of chromium (Cr) in an outer NiCoCrAlY layer.

20. The layer system as claimed in claim 19,

wherein the material of the lower ceramic layer comprises zirconium oxide.

21. The layer system as claimed in claim 19,

wherein the outer ceramic layer comprises zirconium oxide.

22. The layer system as claimed in claim 19,

wherein the substrate comprises a nickel-based or cobalt-based superalloy.

23. The layer system as claimed in claim 19,

wherein the material of the ceramic layers is different.

24. The layer system as claimed in claim 19,

wherein the inner ceramic layer has a form which is at least 10% thinner than the outer ceramic layer.

25. The layer system as claimed in claim 19,

wherein the layer system consists of: the substrate, the two-layered metallic bonding layer, the inner ceramic layer, and the outer ceramic layer.

26. The layer system as claimed in claim 19,

wherein the content of cobalt (Co) in the lower NiCoCrAlY layer is the same as or comparable with the content of cobalt (Co) in the outer NiCoCrAlY layer.

27. The layer system as claimed in claim 19,

wherein the difference in the content of chromium (Cr) in the metallic layers is 3% by weight to 13% by weight.

28. The layer system as claimed in claim 19,

wherein the content of aluminum (Al) in the lower NiCoCrAlY layer is the same as or comparable with the content of aluminum (Al) in the outer NiCoCrAlY layer.

29. The layer system as claimed in claim 19,

wherein the content of yttrium (Y) in the lower NiCoCrAlY layer is the same as or comparable with the content of yttrium (Y) in the outer NiCoCrAlY layer.

30. The layer system as claimed in claim 19,

wherein the lower NiCoCrAlY layer has the following composition in % by weight: cobalt (Co): 24%-26%, chromium (Cr): 11%-15%, aluminum (Al): 10.0%-12.0%, yttrium (Y): 0.2%-0.6%, nickel.

31. The layer system as claimed in claim 19,

wherein the upper NiCoCrAlY layer has the following composition in % by weight: cobalt (Co): 24%-26%, chromium (Cr): 23%-25%, aluminum (Al): 9%-12.0%, yttrium (Y): 0.2%-0.6%, nickel.

32. The layer system as claimed in claim 19,

which has no gradients in the content of chromium (Cr) in the layer.

33. The layer system as claimed in claim 19,

wherein a thermally grown oxide layer is formed or is present on the outer NiCoCrAlY layer.

34. The layer system as claimed in claim 19,

wherein the metallic bonding layer has a total thickness of 180 μm to 300 μm.

35. The layer system as claimed in claim 19,

wherein the powder used for the upper NiCoCrAlY layer is coarser than the grain size of the powder for the lower NiCoCrAlY layer, such that the upper layer comprises larger grains than the lower layer.

36. The layer system as claimed in claim 19,

which comprises no rhenium (Re) in the metallic layers.