WO2015071015A1

WO2015071015A1 - Porous ceramic coating system

Info

Publication number: WO2015071015A1
Application number: PCT/EP2014/069907
Authority: WO
Inventors: Werner Stamm
Original assignee: Siemens Aktiengesellschaft
Priority date: 2013-11-15
Filing date: 2014-09-18
Publication date: 2015-05-21
Also published as: DE102013223327A1

Abstract

The invention relates to a coating system having two layers of porous coatings, comprising a tightly controlled and matched porosity.

Description

Porous ceramic coating system

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

Ceramic protective coatings are often used on high temperature components to protect the metallic substrate from higher temperatures.

In this case, the ceramic layers have a certain porosity, on the one hand to reduce the thermal conductivity and to set a certain ductility.

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

The object is achieved by a layer system according to claim 1. In the subclaims further advantageous measures are listed, which can be combined with each other as desired.

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

a layer system,

a gas turbine,

a turbine blade,

a combustion chamber,

a list of superalloys

The description and figures represent only embodiments of the invention. In Figure 1, the layer system is shown schematically. The layer system 1 is preferably a turbine blade 120, 130 of a turbine, a steam turbine, a gas turbine 100 (FIG. 2) for stationary operation or for aircraft.

Preferably, the substrate 4 comprises a nickel- or cobalt-based superalloy of an alloy of FIG. Preferably, this is a nickel-based superalloy.

On the substrate 4 is preferably a metallic bonding layer 7, in particular of the type MCrAl or MCrAlX present (M = Ni, Co, Fe, preferably Ni, Co). On a metallic bonding layer 7, a ceramic layer 16 is applied, wherein between the metallic

Layer 7 and the ceramic layer, either an oxide layer (TGO) is deliberately generated or applied or forms during the ceramic coating or during the operation of a layer system 1 with 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 <14.5%, especially 11% to 13% (preferably vol%).

The layer thickness of the inner ceramic layer 19 is preferably 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 + 25μπι, while the outer ceramic layer 22 has a thickness of

> 100μπι has. The outer ceramic layer 22 has a porosity of> 15% to 22%, in particular of 16% to 20%, and preferably represents the outermost ceramic layer 22.

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

Preferably, this material is also used for the outer ceramic layer 22, but also a

Pyrochlormaterial can be used as

Gadolinium zirconate.

The selection of porosity of the ceramic layers surprisingly led to a longer service life compared to a dense thick porous layer. Furthermore, according to the invention, a protective layer 7 (FIG. 1) for protecting a component against corrosion and oxidation at a high temperature has substantially the following elements (indication of the parts in% by weight):

nickel

Co: 22% - 24%

Cr: 14% - 16%

AI: 10.5% - 11.5%

0.2% -0.4% rare-earth element (yttrium,...) And / or scanning dium (Sc): in particular yttrium,

optional

Ta: 0.3% - 0.9%.

The list of alloying elements Ni, Co, Cr, Al, Y, Ta is preferably not exhaustive. Nickel preferably forms the matrix.

Preferably, the list of Ni, Co, Cr, Al, Y, Ta is exhaustive. The contents of the alloying elements Co, Cr, Al, Y, Ta have the following advantages: Medium high Co content:

Widening of the beta / gamma field, avoiding brittle phases such as the alpha phases.

Average Cr content:

Sufficiently high to increase activity of AI to A1 ₂ 0 ₃ - formation;

low enough to avoid brittle phases (alpha-chromium or sigma phase). Medium high Al content:

Sufficiently high for AI activity to form a stable Al ₂ O ₃ layer;

low enough to avoid embrittlement effects. Low Y content:

Sufficient high enough to produce enough Y-aluminate to form Y-containing "pegs" with low oxygen contamination;

low enough to negatively accelerate the oxide layer growth of the Al ₂ O ₃ layer.

Tantalum has a positive influence on the phase stability of the γ 'phase or shifts the transition to higher temperatures and thus slows the phase degradation by the consumption of aluminum in the layer.

It should be noted that the proportions of the individual elements are particularly harmonized with regard to their effects, which are to be seen in particular in connection with the element silicon. If the proportions are so dimensioned that no silicon precipitates form, advantageously no brittle phases arise during the use of the protective layer, so that the runtime behavior is improved and prolonged. This is done not only by a low chromium content, but also, taking into account the influence of aluminum on the phase formation, by accurately measuring the content of aluminum.

In interaction with the reduction of the brittle phases, which have a negative effect especially under higher mechanical properties, the reduction of the mechanical stresses by the selected nickel content improves 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 100 (FIG. 3) with a further increase in the inlet temperature. During operation, there is hardly any embrittlement because the layer hardly

Chromium-silicon precipitates, which become brittle in the course of use.

The powders are applied, for example, by plasma spraying (APS, LPPS, VPS, ...) to form a protective layer. Other methods are also conceivable (PVD, CVD, SPPS, ...). The protective layer 7 described also acts as a primer layer to a superalloy.

On this protective layer 7 further layers, in particular ceramic thermal barrier coatings 16 are applied.

In the case of a component 1, the protective layer 7 is advantageously applied to a substrate 4 made of a nickel or cobalt-based superalloy (FIG. 5). Compositions of this type are known as casting alloys under the designations GTD222, IN939, IN6203 and Udimet 500. Further alternatives for the substrate 4 (FIG. 2) of the component 1, 120, 130, 155 are listed in FIG.

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

The protective layer 7 is particularly suitable for protecting the component 1, 120, 130, 155 against corrosion and oxidation, while the component at a material temperature of about 950 ° C, in aircraft turbines also about 1100 ° C, with a

Flue gas is applied.

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

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

The layer system 1 has a substrate 4.

The substrate 4 may be metallic and / or ceramic. Especially with turbine components, such as Turbine Run 120

2 or 3), heat shield elements 155 (FIG. 4) as well as other housing parts of a steam or gas turbine 100 (FIG. 2), the substrate 4 has a nickel-cobalt structure (FIG. 3) or guide vanes 130 (FIGS or iron-based superalloy, in particular it consists of.

Preferably, nickel-based superalloys (Figure 5) are used.

On the substrate 4, the protective layer 7 according to the invention is present.

Preferably, this protective layer 7 is applied by plasma spraying (VPS, LPPS, APS,...). This can be used as outer layer (not shown) or intermediate layer (FIG. 1).

Preferably, the layer system consists of substrate 4,

Protective layer 7 and ceramic thermal barrier coating 16, optionally a TGO under the thermal barrier coating 16th

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

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

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

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

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

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

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

During operation of the gas turbine 100, air 105 is sucked in and compressed by the compressor 105 through the intake housing 104. The compressed air provided at the turbine-side end of the compressor 105 is supplied to the burners 107 where it is mixed with a fuel. The mixture is then burned to form the working medium 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 highest 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 may have a directional structure, i. they are monocrystalline (SX structure) or have only longitudinal grains (DS structure).

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

Such superalloys are 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 vane 130 has a guide vane foot (not shown here) facing the inner housing 138 of the turbine 108 and a vane head opposite the vane foot. The vane head faces the rotor 103 and fixed to a mounting ring 140 of the stator 143. FIG. 3 shows a perspective view of a moving blade 120 or guide blade 130 of a turbomachine that extends along a longitudinal axis 121. The turbomachine may be a gas turbine of an aircraft or a power plant for power generation, a steam turbine or a compressor.

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

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

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

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

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

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

The blade 120, 130 can be made by a casting process, also by directional solidification, by a forging process, by a milling process or combinations thereof. Workpieces with a monocrystalline structure or structures are used as components for machines which are exposed to high mechanical, thermal and / or chemical stresses during operation.

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

Here, dendritic crystals are aligned along the heat flow and form either a columnar grain structure (columnar, i.e., grains that run the full length of the workpiece and here, in common usage, are referred to as directionally solidified) or a monocrystalline structure, i. the whole workpiece consists of a single crystal. In these processes, one must avoid the transition to globulitic (polycrystalline) solidification, since undirected growth inevitably results in transverse and longitudinal grain boundaries, which negate the good properties of the directionally solidified or monocrystalline component.

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

Such methods are known from US Pat. No. 6,024,792 and EP 0 892 090 A1.

Likewise, the blades 120, 130 may have coatings against corrosion or oxidation, e.g. For example, (MCrAlX; 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 the rare earth, or hafnium (Hf)). Such alloys 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 = thermal grown oxide layer) is formed on the MCrAlX layer (as an intermediate layer or as the outermost layer).

Preferably, the layer composition comprises Co-30Ni-28Cr-8A1-0, 6Y-0, 7Si or Co-28Ni-24Cr-10Al-0, 6Y. In addition to these cobalt-based protective coatings, nickel-based protective layers such as Ni-IOCr-12A1-0.6Y-3Re or Ni-12Co-21Cr-IIAl-O, 4Y-2Re or Ni-25Co-17Cr-10A1-0,4Y-1 are also preferably used , 5Re.

On the MCrAlX may still be present a thermal barrier coating, which is preferably the outermost layer, and consists for example of Zr0 ₂ , Y ₂ 0 ₃ -Zr0 ₂ , that is, it is not, partially or completely stabilized by yttria

and / or calcium oxide and / or magnesium oxide.

The thermal barrier coating covers the entire MCrAlX layer.

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

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

MCrAlX layer.

The blade 120, 130 may be hollow or solid. If the blade 120, 130 is to be cooled, it is hollow and may still 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 around a rotation axis 102 open into a common combustion chamber space 154, which produce flames 156 , For this purpose, the combustion chamber 110 is configured in its entirety as an annular structure, which is positioned around the axis of rotation 102 around. To achieve a comparatively high efficiency, the combustion chamber 110 is designed for a comparatively high temperature of the working medium M of about 1000 ° C to 1600 ° C. In order to lend a comparatively long service life to these operating parameters, which are unfavorable for the materials, the combustion chamber wall 153 is provided on its side facing the working medium M with an inner lining formed of heat shield elements 155.

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

Each heat shield element 155 made of an alloy is equipped on the working 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 may be similar to the turbine blades, 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 Si - Lizium 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 B1, EP 0 412 397 B1 or EP 1 306 454 A1. On the MCrAlX, for example, a ceramic thermal barrier coating may be present and consists for example of Zr0 ₂ , Y ₂ 0 ₃ -Zr0 ₂ , ie it is not, partially or completely stabilized by yttrium oxide and / or calcium oxide and / or magnesium oxide.

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

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

Claims

Patent claims

1. Shift system

at least exhibiting

a substrate (4),

a metallic connection layer (7) on the substrate (4),

in particular directly on the substrate (4),

optionally an oxide layer directly on the metallic connection layer (7),

an inner ceramic layer (19) with a porosity, in particular in vol%,

from 8% to <15% on the connection layer (7) or on the

oxide layer,

especially from 9% to <14.5%,

especially from 11% - 13%,

and an outer ceramic layer (22),

in particular an outermost ceramic layer (22) with a porosity of > 15% to 22%,

especially from 16% to 20%,

especially 18%,

on the inner ceramic layer (19),

a NiCoCrAlX layer made of an alloy,

which contains at least the following elements (data in wt.%): 22% - 24% cobalt (Co), in particular 23% cobalt (Co),

14% - 16% chromium (Cr), in particular 15% - 16% chromium (Cr), especially 15% chromium (Cr),

10.5% - 11.5% aluminum (AI), especially llwt% aluminum (AI),

0.2% - 0.4%, in particular 0.3%, of at least one metal from the group comprising scandium (Sc) and / or the rare earth elements, especially yttrium (Y),

optionally 0.3% to 0.9% tantalum (Ta),

in particular 0.4% to 0.7% tantalum (Ta),

especially 0.5% tantalum (Ta),

Nickel (Ni), especially the remainder nickel (Ni).

2. Layer system according to claim 1,

in which the material of the lower ceramic layer (19) is zirconium oxide,

in particular partially stabilized zirconium oxide,

particularly yttrium-partially stabilized zirconium oxide,

in particular consists of it.

Layer system according to one or both of claims 1 or 2,

in which the outer ceramic layer (22) is zirconium oxide, in particular partially stabilized zirconium oxide,

particularly yttrium-partially stabilized zirconium oxide,

in particular consists of it.

4. Layer system according to one or more of claims 1, 2 or 3,

in which the substrate (4) has a nickel or cobalt-based superalloy,

in particular consists of it.

5. Layer system according to one or more of claims 1, 2, 3 or 4,

in which the material of the ceramic layers (19, 22) is different and

in particular the outer ceramic layer (22).

has pyrochlore structure,

especially gadolinium zirconate.

6. Layer system according to one or more of claims 1, 2, 3, 4 or 5,

in which the inner ceramic layer (19) is at least 10%, in particular at least 20%,

is thinner than the outermost ceramic layer (22).

7. Layer system according to one or more of claims 1, 2, 3, 4, 5 or 6,

in which the shift system consists of:

substrate (4),

single-layer metallic connecting layer (7),

optionally oxide layer on the metallic connection layer (7),

inner ceramic layer (19),

outermost ceramic layer (22).

8. Layer system according to claim 1,

containing 0.3% by weight of yttrium (Y) in the alloy.

9. Layer system according to claim 1 or 2,

does not contain rhenium (Re) in the alloy.

10. Layer system according to one or more of the previous claims,

which does not contain silicon (Si) in the alloy.

11. Layer system according to claim 1, 2, 3, 4 or 5,

which contains tantalum (Ta),

in particular contains at least 0.4% by weight of tantalum (Ta), in particular contains 0.5% by weight of tantalum (Ta).

12. Layer system according to one or more of the previous claims,

not containing zirconium (Zr) and/or not containing titanium (Ti) and/or not containing gallium (Ga) and/or not containing germanium (Ge) and/or not containing

Platinum (Pt) and/or not containing hafnium (Hf)

and/or not containing cerium (Ce) and/or not containing iron (Fe) and/or not containing palladium (Pd) and/or

not containing boron (B) and/or

not containing carbon (C).

13. Layer system according to one or more of the previous claims,

with an alloy consisting of cobalt, chromium, aluminum, yttrium, nickel

and optional component tantalum.

14. Layer system according to one or more of the previous claims 1 to 7,

with an alloy consisting of cobalt, chromium, aluminum, yttrium, nickel and tantalum.

15. Layer system according to one or more of claims 1 to 14,

in which a thermally grown oxide layer is formed or is present on the outer NiCoCrAlY layer (13).

16. Layer system according to one or more of claims 1 to 15,

in which the metallic bonding layer (7) has a total thickness of 180μπι to 300μπι.