METHOD FOR THE PRODUCTION AND RESTORATION OF CERAMIC THERMAL INSULATION LAYERS IN GAS TURBINES

Description

[0001] The invention relates to a method for producing gas turbines that are designed to be flexible.

[0002] Gas turbines can be operated in base load mode or, in particular, in peak load mode for electricity generation.

[0003] The requirements for each condition vary.

[0004] An optimized gas turbine configuration that meets both requirements would always represent a compromise. EP 1806430 A1 teaches the reapplication of a new single-layer ceramic thermal protection coating with a porosity of 22-28% to decoated turbine blades for reuse.

[0005] US 2009 / 0324841 A1 teaches the application of a new single-layer or multi-layer ceramic thermal protection coating to decoated turbine blades for reuse.

[0006] WO 2007 / 112713 A1 teaches to provide two or three-layer coatings with a gradually increasing porosity, e.g. inside 9-11%, outside 20-27% and when using a third layer accordingly in between.

[0007] EP 2 407 579 A1 teaches applying a new ceramic thermal protection layer to decoated turbine blades for reuse. A two-layer thermal protection layer is disclosed, which is fixed by means of an adhesive layer applied directly to the substrate. The porosity of the inner layer is 10-18%, and that of the outer layer is >21%.

[0008] US 2002 / 003460 A1 discloses a special coating system for turbine blades with two different porous ceramic (ZrO2) protective layers.

[0009] It is therefore the object of the invention to solve this problem.

[0010] The object is achieved by a method for producing a second gas turbine according to claim 1.

[0011] The subclaims list further advantageous measures which can be combined to achieve further advantages.

[0012] They show: Figures 1 - 3 show embodiments, Figure 4 shows a pore distribution of a ceramic layer, Figure 5 shows a turbine blade and Figure 6 shows a gas turbine.

[0013] The description presents embodiments of the invention as well as embodiments that do not fall under the claims.

[0014] A maintenance interval of gas turbines 100 ( Fig. 6 ) is determined by recording operating hours and starts, which depend on the operating mode and certain factors. Maintenance must be performed whenever the hourly or start limit is reached.

[0015] If, depending on the application of the gas turbine, maintenance is necessary or if the application requires an overhaul or a different application, the configuration of the gas turbine 100 is changed. Definitions of terms:

[0016] First gas turbine has 1st turbine blade with 1st thermal insulation layer.

[0017] Second gas turbine has turbine blades with ceramic thermal insulation coatings a) in which the 1st turbine blades (= second turbine blade) and / or b) new, unused turbine blades (= new second turbine blades) are used and each have a 2nd thermal insulation layer that is clearly distinguishable from the 1st thermal insulation layer.

[0018] If a single-layer thermal insulation layer as described above was previously present in this first gas turbine during operation, a two-layer and / or two-layer ( Fig. 3 ), a thicker one ( Fig. 1) or more porous ceramic thermal insulation layer is used for the turbine blades 120, 130.

[0019] The turbine blades for the second gas turbine can originally be (the same substrate) the first turbine blades of the first gas turbine or other gas turbines that were already in use, were refurbished accordingly and result in second turbine blades through recoating, or they can be new second turbine blades in which newly manufactured (newly cast), not yet used turbine blades are coated differently than the first turbine blades of the first gas turbine.

[0020] It is also possible, if the gas turbine 100 has a two-layer ceramic thermal insulation coating on the turbine blades 120, 130 in base load operation, to apply a single-layer TBC so that it can then be used in peak load operation (daily starter) ( Fig. 2 ).

[0021] For peak load operation, only a single-layer ceramic coating with a uniform porosity is preferably used. For peak load operation, the ceramic thermal barrier coating on the turbine blades 120, 130 preferably has a high porosity of 18% ± 4%.

[0022] In base load operation, however, a two-layer thermal insulation layer 13 is used, or according to the invention, new second ceramic thermal insulation layers 7‴, 10' ( Fig. 3 ).

[0023] Agglomerated, sintered powder is preferably used as starting powder for the ceramic layers 7', 7", 7‴, 10', 13'.

[0024] Each ceramic sprayed layer is applied in coating layers. Two-layer coating means that a second layer differs from a first, underlying layer in terms of porosity, microstructure, and / or chemical composition.

[0025] As the lower layer, a ceramic layer 7 with a porosity of 12% ± 4% is preferably used, which preferably has a layer thickness of 75µm to 150µm.

[0026] Above this, a porosity of 18% ± 4% is injected or is present as the outer ceramic layer 10.

[0027] However, the difference in porosity is at least 2%, especially at least 4%. Porosity variations during production are known. No variations are observed within a batch, i.e., within a set of blades.

[0028] A ceramic layer 7 with a porosity of 18% ± 4% is also preferably used as the lower layer, preferably having a layer thickness of 75 µm to 150 µm. Above this, the outer ceramic layer 10 is sprayed or is present with a porosity of 25% ± 4%.

[0029] However, the difference in porosity is at least 2%, especially at least 4%. Porosity variations during production are known. No variations are observed within a batch, i.e., within a set of blades.

[0030] To detect porosities in ceramic layers or ceramic layers ( Figures 1 - 3 ), coarse grains can be used during spraying and use of polymers or smaller grains with polymer can be used, where coarse means at least 20% larger mean particle diameter.

[0031] A two-layer ceramic layer 7, 10 can be produced using different injection molding processes: the lower layer 7 is injection molded without polymer and the upper layer 10 is injection molded with polymer.

[0032] This results in larger pores in the upper layer 10, ie the average pore diameter d 10 increases compared to the average pore diameter d 7 of the lower layer 7 ( Fig. 4). This is not necessarily the case. Higher porosity is often achieved simply by increasing the number of pores of the same pore size.

[0033] Preferably, the same powder is used, i.e. the same grain size distribution.

[0034] Zirconium oxide (ZrO 2 ) for the ceramic layers of the thermal insulation coatings preferably has a monoclinic content of ≤ 3%, in particular ≤ 1.5%. A ceramic layer or layer 7, 7', 10, 13 ( Figures 1 - 3 ) on the turbine blade 120, 130.

[0035] The minimum proportion of the monoclinic phase is at least 1%, especially 0.5%, in order not to increase the cost of the powder too much.

[0036] By changing the configuration of the first thermal insulation layer 7', 7", 13', a different, second gas turbine is produced, which is optimized for its field of application.

[0037] The Figure 5shows a perspective view of a rotor blade 120 or guide vane 130 of a turbomachine extending along a longitudinal axis 121.

[0038] The turbomachine can be a gas turbine of an aircraft or a power plant for generating electricity, a steam turbine or a compressor.

[0039] The blade 120, 130 has, along the longitudinal axis 121, a fastening region 400, an adjacent blade platform 403, a blade airfoil 406 and a blade tip 415.

[0040] As a guide vane 130, the vane 130 may have a further platform at its vane tip 415 (not shown).

[0041] In the fastening region 400, a blade root 183 is formed, which serves to fasten the rotor blades 120, 130 to a shaft or a disk (not shown).

[0042] The blade base 183, for example, is designed as a hammer head. Other designs such as a fir tree or dovetail base are possible.

[0043] The blade 120, 130 has a leading edge 409 and a trailing edge 412 for a medium flowing past the blade 406.

[0044] In conventional blades 120, 130, for example, solid metallic materials, in particular superalloys, are used in all areas 400, 403, 406 of the blade 120, 130.

[0045] 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.

[0046] The blade 120, 130 can be manufactured by a casting process, also by means of directional solidification, by a forging process, by a milling process or combinations thereof.

[0047] Workpieces with a single-crystal structure or structures are used as components for machines that are exposed to high mechanical, thermal and / or chemical stresses during operation.

[0048] The production of such monocrystalline workpieces is achieved, for example, by directional solidification from the melt. These are casting processes in which the liquid metallic alloy solidifies into a monocrystalline structure, i.e., the monocrystalline workpiece, or solidifies in a directionally solidified manner.

[0049] In this process, dendritic crystals are aligned along the heat flow and form either a columnar grain structure (i.e., grains that run the entire length of the workpiece and are commonly referred to as directionally solidified) or a monocrystalline structure, i.e., the entire workpiece consists of a single crystal. In these processes, the transition to globulitic (polycrystalline) solidification must be avoided, since undirected growth inevitably leads to the formation of transverse and longitudinal grain boundaries, which negate the beneficial properties of the directionally solidified or monocrystalline component.

[0050] When we generally speak of directionally solidified structures, we mean both single crystals that have no grain boundaries or at most small-angle grain boundaries, and columnar crystal structures that have longitudinal grain boundaries but no transverse grain boundaries. These latter crystalline structures are also referred to as directionally solidified structures.

[0051] Such processes are known from US-PS 6,024,792 and EP 0 892 090 A1.

[0052] Likewise, the blades 120, 130 may have coatings against corrosion or oxidation, e.g., (MCrAlX; M is at least one element from the group consisting of iron (Fe), cobalt (Co), and nickel (Ni), X is an active element and represents yttrium (Y) and / or silicon and / or at least one rare earth element, 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.

[0053] The density is preferably 95% of the theoretical density.

[0054] A protective aluminum oxide layer (TGO = thermal grown oxide layer) forms on the MCrAlX layer (as an intermediate layer or as an outermost layer).

[0055] The coating composition preferably comprises Co-30Ni-28Cr-8Al-0.6Y-0.7Si or Co-28Ni-24Cr-10Al-0.6Y. In addition to these cobalt-based protective coatings, nickel-based protective coatings 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 are also preferably used.

[0056] A thermal barrier layer may also be present on the MCrAlX, 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 completely stabilized by yttrium oxide and / or calcium oxide and / or magnesium oxide.

[0057] The thermal barrier coating covers the entire MCrAlX layer. Suitable coating processes, such as electron beam evaporation (EB-PVD), create columnar grains in the thermal barrier coating.

[0058] Other coating processes are conceivable, such as atmospheric plasma spraying (APS), LPPS, VPS, or CVD. The thermal barrier coating can contain porous grains with micro- or macrocracks for improved thermal shock resistance. The thermal barrier coating is therefore preferably more porous than the MCrAlX coating.

[0059] Refurbishment means that components 120, 130 may need to be freed of protective coatings after use (e.g., by sandblasting). This is followed by the removal of corrosion and / or oxidation layers or products. If necessary, cracks in components 120, 130 are also repaired. The components 120, 130 are then recoated and reused.

[0060] The blade 120, 130 can be hollow or solid. If the blade 120, 130 is to be cooled, it is hollow and may also have film cooling holes 418 (indicated by dashed lines).

[0061] The Figure 6 shows an example of a gas turbine 100 in a longitudinal section.

[0062] The gas turbine 100 has inside a rotor 103 with a shaft 101, which is rotatably mounted about a rotation axis 102 and is also referred to as a turbine rotor.

[0063] Along the rotor 103 there follow an intake housing 104, a compressor 105, a toroidal combustion chamber 110, in particular an annular combustion chamber, with several coaxially arranged burners 107, a turbine 108 and the exhaust housing 109.

[0064] The annular combustion chamber 110 communicates with an annular hot gas duct 111, for example. There, four turbine stages 112 connected in series form the turbine 108.

[0065] Each turbine stage 112 is formed, for example, from two blade rings. Viewed in the flow direction of a working medium 113, a row of guide vanes 115 is followed in the hot gas duct 111 by a row 125 formed from rotor blades 120.

[0066] The guide vanes 130 are attached to an inner housing 138 of a stator 143, whereas the rotor blades 120 of a row 125 are attached to the rotor 103, for example by means of a turbine disk 133.

[0067] A generator or a working machine (not shown) is coupled to the rotor 103.

[0068] During operation of the gas turbine 100, air 135 is drawn in by the compressor 105 through the intake casing 104 and compressed. The compressed air provided at the turbine end of the compressor 105 is fed to the burners 107, where it is mixed with a fuel. The mixture is then combusted in the combustion chamber 110 to form the working medium 113. From there, the working medium 113 flows along the hot gas duct 111 past the guide vanes 130 and the rotor blades 120. At the rotor blades 120, the working medium 113 expands, transmitting momentum, so that the rotor blades 120 drive the rotor 103, which drives the driven machine coupled to it.

[0069] 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, as seen in the flow direction of the working medium 113, are subjected to the greatest thermal loads, along with the heat shield elements lining the annular combustion chamber 110.

[0070] In order to withstand the temperatures prevailing there, they can be cooled using a coolant.

[0071] Likewise, substrates of the components can have a directional structure, i.e. they are single-crystalline (SX structure) or have only longitudinally oriented grains (DS structure).

[0072] For example, iron-, nickel- or cobalt-based superalloys are used as materials for the components, in particular for the turbine blades 120, 130 and components of the combustion chamber 110.

[0073] 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.

[0074] Likewise, the blades 120, 130 can have anti-corrosive coatings (MCrAlX; M is at least one element from the group consisting of iron (Fe), cobalt (Co), and 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). 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.

[0075] A thermal insulation layer may be present on the MCrAlX, and may consist, for example, of ZrO 2 , Y 2 O 3 -ZrO 2 , ie it is not, partially or completely stabilized by yttrium oxide and / or calcium oxide and / or magnesium oxide.

[0076] Using suitable coating processes such as electron beam evaporation (EB-PVD), stem-shaped grains are created in the thermal barrier coating.

[0077] The guide vane 130 has a guide vane root (not shown here) facing the inner casing 138 of the turbine 108 and a guide vane tip opposite the guide vane root. The guide vane tip faces the rotor 103 and is secured to a mounting ring 140 of the stator 143.

Claims

1. Process for producing a second gas turbine, in which, proceeding from a first gas turbine, a first ceramic thermal barrier coating (7) is removed from first turbine blades or vanes (120, 130) of the first gas turbine and new, second ceramic thermal barrier coatings (7‴, 10') are applied to the first turbine blades or vanes (120, 130) from which the coating has been removed to produce second turbine blades or vanes, characterized in that the new, second ceramic thermal barrier coatings (7‴, 10') differ significantly from the first ceramic thermal barrier coating (7), in which case a difference means that the layer nature is different due to the single-layer or two-layer nature of the ceramic thermal barrier coatings (7‴, 10'), and wherein the second turbine blades or vanes are incorporated in the second gas turbine.

2. Process according to Claim 1, in which a single-layer thermal barrier coating (7) is removed from the first turbine blades or vanes.

3. Process according to Claim 1 or 2, in which the porosity of the second ceramic thermal barrier coatings (7‴, 10') of the second turbine blades or vanes is elevated compared to the porosity of the thermal barrier coating of the first turbine blades or vanes.

4. Process according to Claim 1 or 2, in which the porosity of the second ceramic thermal barrier coatings (7"', 10') of the second turbine blades or vanes is lowered compared to the porosity of the thermal barrier coating of the first turbine blades or vanes.

5. Process according to one or more of Claims 1 to 3, in which a thinner ceramic thermal barrier coating (7) as the first ceramic thermal barrier coating is replaced by an overall thicker ceramic thermal barrier coating (13') formed from the second ceramic thermal barrier coatings (7‴, 10') of the second turbine blades or vanes, the difference in the thickness being at least +50 µm.

6. Process according to one or more of Claims 1 to 4, in which a thicker ceramic thermal barrier coating as the first ceramic thermal barrier coating is replaced by an overall thinner ceramic thermal barrier coating (13') formed from the second ceramic thermal barrier coatings (7"', 10') of the second turbine blades or vanes, the difference in the thickness being at least -50 µm.

7. Process according to one or more of Claims 2 to 6, in which the second ceramic thermal barrier coatings are produced with a bottommost ceramic layer (7‴) having a porosity of 12% ± 4% and with an outer ceramic layer (10') having a porosity of 18% ± 4%, the absolute difference in the porosity of the ceramic layers (7‴, 10') being at least 2%, in particular at most 4%.

8. Process according to one or more of Claims 2 to 6, in which the second ceramic thermal barrier coatings are produced with a bottommost ceramic layer (7‴) having a porosity of 18% ± 4% and with an outer ceramic layer (10') having an identical porosity of 18% ± 4%.

9. Process according to one or more of Claims 2 to 6, in which the second ceramic thermal barrier coatings are produced with a bottommost ceramic layer (7‴) having a porosity of 18% ± 4% and with an outer ceramic layer (10') having a porosity of 25% ± 4%, the absolute difference in the porosity of the ceramic layers (7‴, 10') being at least 2%, in particular at most 4%.

10. Process according to one or more of Claims 2 to 9, in which the bottom layer (7‴) of the second ceramic thermal barrier coatings is designed to be thinner, in particular at least 20% thinner, than the top layer (10'), in which the bottom layer (7‴) of the second ceramic thermal barrier coatings has a thickness of 75 µm to 150 µm, the total coating thickness of the second ceramic thermal barrier coatings being 500 µm to 800 µm.

11. Process according to one or more of Claims 2 to 10, in which partially stabilized zirconium oxide is used for the bottom ceramic layer (7‴) and partially stabilized zirconium oxide is used for the top ceramic layer (10').

12. Process according to one or more of the preceding claims, in which zirconium oxide is used for the ceramic thermal barrier coatings (7‴, 10') , and the monoclinic proportion of the powder to be sprayed is less than 3%, and is at least 0.3%.

13. Process according to one or both of Claims 11 and 12, in which the tetragonal proportion has the greatest proportion of at least 60% in zirconium oxide.

14. Process according to Claim 12 or 13, in which a heat treatment reduces the monoclinic proportion of the zirconium oxide below the detection limit.

15. Process according to one or more of Claims 2 to 14, in which the bottom layer (7‴) is sprayed without polymer and the top layer (10') is sprayed with polymer.

16. Process according to one or more of Claims 2 to 15, in which the mean pore diameter (d10) of the top ceramic layer (10') is produced so as to be greater than the mean pore diameter (d7) of the bottom ceramic layer (7‴), in particular by at least 20 µm.

17. Process according to one or more of Claims 2 to 16, in which the same powder with the same composition and with the same grain size distribution is used for the ceramic layers (7‴, 10').

18. Process according to Claims 2 to 16, in which a different material is used for the bottom ceramic layer (7‴) than for the top ceramic layer (10'), in particular zirconium oxide for the bottom layer (7‴), very particularly a pyrochlore for the top layer (10').