Method for treating a gas turbine blade and gas turbine having said blade

Abstract

The use of different ceramic layers allows different configurations of gas turbines to be produced each of which is optimized for a respective use of base load operation or peak load operation.

Description

TECHNICAL FIELD

The invention relates to a process for producing gas turbines of flexible design, to gas turbines and to methods for operating gas turbines.

TECHNICAL BACKGROUND

For generating electricity, gas turbines can be operated in base load operation or in particular in peak load operation.

The demands for satisfaction of these respective conditions are different. An optimized configuration of the gas turbine which satisfies both demands would always represent a compromise. It is therefore an object of the invention to solve this problem.

BRIEF DESCRIPTION OF THE DRAWINGS

FIGS. 1-3 show an exemplary embodiment of the invention,

FIG. 4 shows a pore distribution of a ceramic coating,

FIG. 5 shows a turbine blade or vane, and

FIG. 6 shows a gas turbine.

DESCRIPTION OF AN EMBODIMENT

The description represents merely an exemplary embodiment of the invention.

A maintenance interval for gas turbine 100 (which is shown in FIG. 6) is determined by ascertaining the operational hours and starts, and these are dependent on the mode of operation and specific factors. The maintenance is to be carried out in each case when the hour or start limit has been reached.

Depending on the field of use of the gas turbine, if it is then necessary to carry out maintenance or if the use demands reconditioning or another use beforehand, the configuration of the gas turbine 100 is altered.

Definitions of the Terms

First gas turbine has first turbine blade or vane with first thermal barrier coating.

Second gas turbine has turbine blades or vanes with ceramic thermal barrier coatings,

a) in which the first turbine blades or vanes (=second turbine blade or vane) and/or

b) new, unconsumed turbine blades or vanes (=new, second turbine blades or vanes)

are used, and in each case have a second thermal barrier coating which can be clearly distinguished from the first thermal barrier coating.

If a single-layer thermal barrier coating was present in operation beforehand in said first gas turbine, as described above, a two-layer (FIG. 3), a thicker (FIG. 1) or a more porous ceramic thermal barrier coating is used for the turbine blades or vanes 120, 130 for the renewed use in base load operation.

The origin (that is, the same substrate) of the turbine blades or vanes for the second gas turbine may be the first turbine blades or vanes of the first gas turbine or other gas turbines, which were already in use, have been appropriately refurbished and then give rise to second turbine blades or vanes through recoating, or can be new, second turbine blades or vanes, in which newly produced (newly cast) turbine blades or vanes which have not yet been used are coated differently to the first turbine blades or vanes of the first gas turbine.

Similarly, it is possible, if the gas turbine 100 had a two-layer ceramic thermal barrier coating on the turbine blades or vanes 120, 130 in base load operation, to apply a single-layer TBC, such that it can then be used in peak load operation (daily starter) (FIG. 2).

For peak load operation, it is preferable to use only a single-layer ceramic coating with a uniform porosity. For peak load operation, the ceramic thermal barrier coating on the turbine blades or vanes 120, 130 preferably has a high porosity of 18%±4%.

In base load operation (base loader), however, a two-layer thermal barrier coating 13 is used (FIG. 3).

It is preferable to use agglomerated, sintered powder as starting powder for the ceramic coatings 7′, 7″, 7′″, 10′, 13′.

Each ceramic sprayed coating is applied in coating layers. Two-layer nature means, however, that a second layer differs from a first, underlying layer in terms of porosity and/or microstructure and/or chemical composition.

A ceramic layer 7 with a porosity of 12%±4% which preferably has a coating thickness of 75 μm to 150 μm is preferably used as the bottom layer.

A layer with a porosity of 25%±4% is sprayed or is present thereabove as the outer ceramic layer 10.

The difference in the porosity is, however, at least 2%, in particular at least 4%. Variations in the porosity during production are known. No variations are to be recorded within a charge, i.e. a blade or vane set.

In order to generate porosities in ceramic coatings or ceramic layers (FIGS. 1-3), the spraying can involve the use of coarse grains and use can be made of polymers or smaller grains with polymer, coarse meaning an at least 20% greater mean particle diameter.

A two-layer ceramic coating 7, 10 can be produced using different spraying processes: the bottom layer 7 is sprayed without polymer and the top layer 10 is sprayed with polymer.

This gives rise to larger pores in the top layer 10, i.e. the mean pore diameter d₁₀ increases compared to the mean pore diameter d₇ of the bottom layer 7 (FIG. 4). This is not necessarily the case. A higher porosity is often only achieved by a higher number of pores of the same pore size.

It is preferable that the same powder is used in this case, i.e. also an identical grain size distribution.

Zirconium oxide (ZrO₂) for the ceramic layers of the thermal barrier coatings preferably has a monoclinic proportion of ≤3%, in particular 1.5%. A ceramic layer or coating 7, 7′, 10, 13 (FIGS. 1-3) on the turbine blade or vane 120, 130 then has corresponding proportions.

The minimum proportion of the monoclinic phase is at least 1%, in particular 0.5%, so as not to excessively increase the costs of the powder.

The change in the configuration of the first thermal barrier coating 7′, 7″, 13′ virtually produces another, second gas turbine optimized for its field of use.

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.

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.

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, for example 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.

Refurbishment means that after they have been used, protective layers may have to be removed from components 120, 130 (e.g. by sand-blasting). Then, the corrosion and/or oxidation layers and products are removed. If appropriate, cracks in the component 120, 130 are also repaired. This is followed by recoating of the component 120, 130, after which the component 120, 130 can be reused.

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. 6 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 101 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 blades or vanes 120, 130 may likewise have coatings protecting against corrosion (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, scandium (Sc) and/or at least one rare earth element, or hafnium). 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 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). 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.

Claims

1. A method for operating a gas turbine system, comprising: altering turbine blades or vanes of a first gas turbine to produce a second gas turbine, the first turbine blades or vanes having first ceramic thermal barrier coatings configured for base load operation;removing completely the first ceramic thermal barrier coating of a first turbine blade or vane of the first gas turbine;applying a new, second ceramic thermal barrier coating configured for peak load operation, different from the base load operation, to the first turbine blade or vane from which the first ceramic thermal barrier coating has been completely removed to produce a second turbine blade or vane by coating with a ceramic and a polymer or with a ceramic with grains of at least 20% greater mean particle diameter to generate pores in the new, second ceramic thermal barrier coating; andstarting the second gas turbine daily instead of continuously operating the second gas turbine;wherein the new, second ceramic thermal barrier coating differs significantly from the completely removed first ceramic thermal barrier coating, in that porosities of the first and the second ceramic thermal barrier coatings are different, and the absolute difference in the reduced or increased porosity is at least 2%, and difference in coating thicknesses of the first and the second ceramic thermal barrier coatings is at least 50 μm;wherein the porosity of the completely removed first ceramic thermal barrier coating, and the porosity of the second ceramic thermal barrier coating are in the range 12%±4% to 25%±4%, andwherein the second turbine blade or vane is incorporated in the second gas turbine.

2. The method as claimed in claim 1, wherein the completely removed first ceramic thermal barrier coating is a two-layer ceramic thermal barrier coating; with a bottommost layer and an outer layer, the bottommost layer having a porosity in the range 25%±4% and the outer layer having a porosity in the range 12%±4%, and the second ceramic thermal barrier coating is a single layer thermal barrier coating with a porosity in the range 18%±4%.

3. The method as claimed in claim 1, wherein the porosity of the second ceramic thermal barrier coating is elevated compared to the porosity of the first ceramic thermal barrier coating.

4. The method as claimed in claim 1, wherein the porosity of the second ceramic thermal barrier coating is lower than the porosity of the first ceramic thermal barrier coating.

5. The method as claimed in claim 1, wherein the first ceramic thermal barrier coating is thinner than the second ceramic thermal barrier coating and the difference in the thicknesses is at least +50 μm.

6. The method as claimed in claim 1, wherein the first ceramic thermal barrier coating is thicker than the second ceramic thermal barrier coating and the difference in the thickness being at least −50 μm.

7. The method as claimed in claim 1, wherein the second ceramic thermal barrier coating is a single layer coating with a porosity in the range 18%±4%, and the first ceramic thermal barrier coating has a bottommost ceramic layer with a porosity of 12%±4% and an outer ceramic layer with a porosity of 25%±4%, wherein an absolute difference in the porosities of the ceramic layers is at least 2%.

8. The method as claimed in claim 7, wherein the mean pore diameter of the outer ceramic layer is greater than the mean pore diameter of the bottommost ceramic layer.

9. The method as claimed in claim 1, wherein the second turbine blade or vane has a ceramic thermal barrier coating consisting of a single layer with a uniform porosity of 18%±4%.

10. The method as claimed in claim 1, wherein the first turbine blade or vane has a two-layer ceramic thermal barrier coating which consists of a bottommost ceramic layer having a porosity of 12%±4% and a top ceramic layer having a porosity of 25%±4%, and the absolute difference in the porosities of the ceramic layers is at least 2%.

11. The method as claimed in claim 10, wherein the bottommost ceramic layer comprises partially stabilized zirconium oxide and the top ceramic layer comprises partially stabilized zirconium oxide.

12. The method as claimed in claim 10, wherein the top ceramic layer has a perovskite or pyrochlore structure, and the bottommost ceramic layer comprises zirconium oxide.

13. The method as claimed in claim 10, wherein the mean pore diameter of the top ceramic layer is greater than the mean pore diameter of the bottommost ceramic layer, by at least 20 μm.