POROUS CERAMIC COATING SYSTEM
A coating system having two layers of porous coatings is provided. The layers include a tightly controlled and matched porosity. The layer system also includes a substrate with a bonding layer. The inner ceramic layer and the outer ceramic layer are formed on the bonding layer. The bonding layer may comprise a MCrAlX alloy.
This application is the US National Stage of International Application No. PCT/EP2011/061320, filed Jul. 5, 2011 and claims the benefit thereof. The International Application claims the benefits of European Patent Office application No. 10007240.4 EP filed Jul. 14, 2010. All of the applications are incorporated by reference herein in their entirety.
FIELD OF INVENTIONThe invention relates to a layer system having two different porous ceramic layers.
BACKGROUND OF INVENTIONCeramic 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 INVENTIONIt is an object of the invention to optimize the thermal and the mechanical properties.
The object is achieved by a layer system as claimed in the claims
The dependent claims list further advantageous measures which can be combined with one another, as desired.
The description and the figures show merely exemplary embodiments of the invention.
DETAILED DESCRIPTION OF INVENTIONIt is preferable for the substrate 4 to comprise a nickel-based or cobalt-based superalloy made of an alloy shown in
A metallic bonding layer, in particular of the MCrAl or MCrAlX type (M═Ni, Co, Fe, preferably Ni, Co), is preferably present on the substrate 4.
Similarly, there may be diffusion layers within the substrate 4, to which a ceramic layer 16 can be applied.
A ceramic layer 16 is applied to the metallic layer 7 or to the substrate 4, wherein an oxide layer (TGO) is either produced deliberately on or applied to the interface, this forming 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 10, 13. The lower ceramic layer 10 has a lower porosity than the outer ceramic layer 13. The porosity of the lower ceramic layer 10 is 8% to 14%, in particular 11% to 12% (preferably % by volume).
It is preferable for the layer thickness of the inner ceramic layer 10 to be at least 10%, in particular 20%, very particularly 50%, thinner than that of the outer ceramic layer 13. The lower layer has a thickness of 100±25 μm, whereas the outer layer has a thickness of >100 μm.
The outer ceramic layer 13 has a porosity of 14% to 18% and is preferably the outermost layer, which is exposed directly to the hot gas.
The material of the lower ceramic layer 10 is partially stabilized, in particular by yttrium, zirconium oxide. This material is preferably also used for the outer ceramic layer 13, although it is also possible to use a pyrochlore material.
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.
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, where, by way of example, four successive turbine stages 112 form the turbine 108.
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.
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, to be present on the MCrAlX, 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.
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).
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 to be present on the MCrAlX, 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.
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
1-8. (canceled)
9. A layer system, comprising:
- a substrate;
- a metallic bonding layer on the substrate;
- an oxide layer on the bonding layer or on the substrate;
- an inner ceramic layer having a porosity of 11% to 12%; and
- an outermost ceramic layer, having a porosity of 16% to 18%, on the inner ceramic layer.
10. The layer system as claimed in claim 9, wherein the metallic bonding layer comprises an MCrAl or MCrAlX alloy.
11. The layer system as claimed in claim 9, wherein the metallic bonding layer consists of an MCrAl or MCrAlX alloy.
12. The layer system as claimed in claim 9, wherein the material of the inner ceramic layer comprises zirconium oxide.
13. The layer system as claimed in claim 9, wherein the material of the inner ceramic layer comprises partially stabilized zirconium oxide.
14. The layer system as claimed in claim 9, wherein the material of the inner ceramic layer comprises zirconium oxide partially stabilized by yttrium.
15. The layer system as claimed in claim 9, wherein the material of the inner ceramic layer consists of zirconium oxide partially stabilized by yttrium.
16. The layer system as claimed in claim 9, wherein the outer layer comprises zirconium oxide.
17. The layer system as claimed in claim 9, wherein the outer layer comprises partially stabilized zirconium oxide.
18. The layer system as claimed in claim 9, wherein the outer layer comprises zirconium oxide partially stabilized by yttrium.
19. The layer system as claimed in claim 9, wherein the outer layer consists of zirconium oxide partially stabilized by yttrium.
20. The layer system as claimed in claim 9, wherein the substrate comprises a nickel-based or cobalt-based superalloy.
21. The layer system as claimed in claim 9, wherein the substrate consists of a nickel-based or cobalt-based superalloy.
22. The layer system as claimed in claim 9, wherein the material of the inner ceramic layer and the material of the outer ceramic layer are different.
23. The layer system as claimed in claim 22, wherein the outer ceramic layer comprises a pyrochlore structure.
24. The layer system as claimed in claim 9, wherein the inner ceramic layer is at least 10% thinner than the outermost ceramic layer.
25. A layer system, consisting of:
- a substrate;
- a metallic bonding layer;
- an oxide layer on the bonding layer;
- an inner ceramic layer; and
- an outermost ceramic layer.
Type: Application
Filed: Jul 5, 2011
Publication Date: May 9, 2013
Inventor: Werner Stamm (Mulheim an der Ruhr)
Application Number: 13/808,259
International Classification: F01D 5/28 (20060101);
