Method for producing a thermal barrier system on a metal substrate of a turbo engine part

Abstract

Method for producing a thermal barrier system on a metal substrate (1) of a turbo engine part, such as a high-pressure turbine blade, the thermal barrier system comprising at least one columnar ceramic layer (31, . . . , 3i, . . . , 3n), characterised in that the method comprises a step of compressing at least one of said at least one columnar ceramic layer (31, . . . 3i, . . . , 3n).

Claims

1. A method for producing a thermal barrier system on a metal substrate of a turbine engine part, said thermal barrier system comprising a stack of a plurality of columnar ceramic layers which have an upper ceramic layer and at least one intermediate ceramic layer arranged under said upper ceramic layer, said upper ceramic layer forming a top of said stack of plurality of columnar ceramic layers, said upper ceramic layer comprising a lower surface which is in contact with said at least one intermediate ceramic layer and an upper surface which is not in contact with any layer, said at least one intermediate ceramic layer being arranged between said upper ceramic layer and said metal substrate, the method comprising the steps of: successively depositing said plurality of columnar ceramic layers on said metal substrate, stacking said at least one intermediate ceramic layer on said metal substrate, compressing said at least one intermediate ceramic layer to tighten spaces between columns of said at least one intermediate ceramic layer, and stacking said upper ceramic layer on said at least one intermediate ceramic layer.

2. The method according to claim 1, wherein the compression is a shot peening, a microbeads peening or a compression by laser shock peening.

3. The method according to claim 2, wherein the compression said at least one intermediate ceramic layer is a shot peening and in that the Almen intensity of said shot peening is between F10A and F42A.

4. The method according to claim 1, wherein said substrate is a nickel or cobalt-based superalloy substrate.

5. The method according to claim 1, wherein at least one of said plurality columnar ceramic layers is a layer of yttriated zirconia.

6. The method according to claim 1, wherein said plurality columnar ceramic layers are obtained by physical vapour deposition.

7. The method according to claim 6, wherein the vapour deposition is an electron beam physical vapour deposition (EBPVD).

8. The method according to claim 1, wherein the method comprises the compression of said upper ceramic layer.

9. The method according to claim 1, wherein the thermal barrier system further comprises a bonding layer arranged between said metal substrate and said plurality columnar ceramic layers.

10. The method according to claim 9, wherein said bonding layer is a layer of an aluminium-forming material comprising an alumina layer on the surface.

11. The method according to claim 9, wherein the method comprises a step of compressing said bonding layer.

12. The method according to claim 11, wherein the compression of said bonding layer is a shot peening and in that the Almen intensity of said shot peening is between F9N and F30A.

13. The method according to claim 1, wherein said turbine engine part is a high pressure turbine blade.

14. The method according to claim 1, wherein the compression of one or more of said at least one intermediate ceramic layers is partial or total.

15. The method according to claim 1, wherein said at least one intermediate ceramic layer is located directly under said upper ceramic layer.

16. The method according to claim 1, wherein the method comprises a step of compressing said upper ceramic layer to tighten spaces between columns of said upper ceramic layer.

17. The method according to claim 9, wherein said bonding layer is sandwiched between said metal substrate and said at least one intermediate ceramic layer.

18. The method according to claim 9, wherein said stack of plurality of columnar ceramic layers comprise also a lower ceramic layer which rests on said bonding layer, said lower ceramic layer being arranged under said at least one intermediate ceramic layer.

19. The method according to claim 18, wherein the method comprises a step of compressing said lower ceramic layer to tighten spaces between columns of said lower ceramic layer.

Description

DESCRIPTION OF THE FIGURES

(1) The invention will be better understood and other details, characteristics and advantages of the invention will appear upon reading the following description provided as a non-limiting example and in reference to the appended drawings, wherein:

(2) FIG. 1 is a schematic, cross-sectional view of a thermal barrier system of a state of the art turbo engine blade,

(3) FIGS. 2 and 3 are schematic, cross-sectional views of a thermal barrier system produced according to a method according to the invention, according to two alternatives of a first embodiment;

(4) FIG. 4 is a schematic, cross-sectional view of a thermal barrier system produced according to a method according to the invention, according to a second embodiment;

(5) FIGS. 5 and 6 are schematic, cross-sectional views of a thermal barrier system produced according to a method according to the invention, according to two alternatives of a third embodiment;

(6) FIG. 7 is a schematic, cross-sectional view of a thermal barrier system produced according to a method according to the invention, according to a fourth embodiment;

(7) FIG. 8 is a schematic, cross-sectional view of a thermal barrier system produced according to a method according to the invention, according to a fifth embodiment.

DETAILED DESCRIPTION

(8) FIG. 1 shows a cross-sectional view of the composition of a thermal barrier system arranged on the surface of a turbine blade, with the latter being bathed by a flow of hot gas represented by an arrow directed towards the left of the figure. The metal that forms the blade, typically a nickel or cobalt-based superalloy, forms a substrate 1 on which is deposited a sublayer made of aluminium 2, referred to as bonding layer, sandwiched between the substrate 1 and a ceramic layer 3. The function of the bonding layer 2 is to retain the ceramic layer 3 and to offer a certain elasticity to the whole in order to make it possible for it to absorb the difference in dilatation, represented by two arrows in the opposite direction, existing between the substrate 1 with high dilatation and the ceramic 3 with low dilatation.

(9) The bonding layer 2 can be of the MCrAlY formula, wherein M designates Fe, Ni, Co and mixtures thereof. It can be obtained by conventional plasma spraying, for example of the APS (Air Plasma Spraying) type. The bonding layer 2 of the MCrAlY type can be replaced with a nickel aluminide or modified with platinum, or with a layer of the gamma/gamma-MCrAlY type.

(10) The ceramic 3 shown here has a columnar structure, which enables lateral movements, due to the appearance of cracks between the columns, and which provides it with a good service life. The aluminium is then put into contact with the oxygen conveyed by the gases that circulate in the stream of the turbo engine, which results in a mediocre thermal conductivity of the barrier and progressive damage to the latter.

(11) The ceramic coating can be formed from a stack of one or more layers, produced for example by an electron beam physical vapour deposition (EBPVD). The first ceramic layer is preferably with a yttriated zirconia base that is partially stabilised (YSZ). For the other ceramic layers, different types of layers can be considered: a mono-oxide, such as for example Al.sub.2O.sub.3 or Y.sub.2O.sub.3, a zirconia doped with one or more rare-earth oxides, a rare-earth zirconate, such as for example Gd.sub.2Zr.sub.2O.sub.7, Sm.sub.2Zr.sub.2O.sub.7 or Yb.sub.4Zr.sub.3O.sub.12, a perovskite, such as for example Ba(Mg.sub.1/3Ta.sub.2/3)O.sub.3, La(Al.sub.1/4Mg.sub.1/2Ta.sub.1/4)O.sub.3, a hexaaluminate, for example of the general formula REMAl.sub.11O.sub.19, wherein RE designates an element ranging from La to Gd in the periodic table, and M designates an element chosen from Mg, Mn to Zn, Cr and Sm, lanthanide orthophosphates.

(12) The thermal barrier system functions to prolong the service life of the blade and to increase the temperature of the gases, and therefore the output of the engine. In service, the structure and the composition of the various constituents of the system change under the action of the sintering of the ceramic layer, of the oxidation of the bonding layer and of the interdiffusion phenomena with the substrate, with consequently a modification in the properties of the various layers and an alteration of the resistance of the interfacial zone. These modifications, associated with the external thermo-mechanical stresses, are at the origin of the roughness of the bonding layer leading to delaminations at the bonding/ceramic layer interface, and, in the end, to the flaking of the thermal barrier system. These degradation processes can be accelerated by the interactions with the external environment.

(13) To overcome this, and according to the invention, the compression of at least one columnar ceramic layer is carried out.

(14) In a first embodiment, the compression of the upper columnar ceramic layer is carried out. Such as shown in FIG. 2, the ceramic coating comprises a single ceramic layer 3, for example of the YSZ type. The ceramic layer 3 undergoes a compression operation C3, so as to tighten the intercolumnar spaces at the surface, which have the effect of: a limitation of the infiltration of the CMAS oxides, an increase in the service life of the thermal barrier system, an improvement in the mechanical properties, such as the surface hardness, an increase in the resistance to erosion, and an increase in the tenacity of the thermal barrier system.

(15) The compression of the ceramic layer 3 is symbolised in FIG. 2 by the reference C3 showing a compressed layer portion. The layer 3 can be compressed partially or entirely, i.e. over all or a portion of the height of the layer 3.

(16) In the alternative shown in FIG. 3, the ceramic coating comprises a plurality of n ceramic layers. A lower layer 31 rests on the bonding layer 2. In the direction of the surface of the thermal barrier system, there is an intermediate layer 3i and the upper layer 3n. The compression of the upper ceramic layer 3n is symbolised in FIG. 2 by the reference C3n. The layer 3n can be compressed partially or entirely, i.e. over all or a portion of the height of the layer 3n. The compression C3n makes it possible to tighten the intercolumnar spaces at the surface of the thermal barrier system and achieves the same advantages as those mentioned for FIG. 2.

(17) In a second embodiment, shown in FIG. 4, each layer of the ceramic coating that comprises n layers is subjected to a compression, partially or entirely. Thus, the first layer 31, for example of the YSZ type, is subjected to a compression C31, each intermediate layer 3i is subjected to a compression C3i and the upper layer 3n is subjected to a compression C3n.

(18) The substrate 1 of the turbine blade is covered beforehand or not with a bonding layer 2 of the MCrAlY type, M designating Fe, Ni, Co and mixtures thereof. The bonding layer 2 can be obtained by conventional plasma spraying, for example of the APS (Air Plasma Spraying) type. The bonding layer 2 of the MCrAlY type can be replaced with a nickel aluminide or modified with platinum, or with a layer of the gamma/gamma-MCrAlY type.

(19) The ceramic coating is formed from a stack of n layers 31, . . . , 3i, . . . , 3n, produced by the electron beam physical vapour deposition (EBPVD) method. The first layer 3i is preferably an yttriated zirconia base that is partially stabilised.

(20) After the production of each ceramic layer by EBPVD, a compression operation is carried out that makes it possible to obtain a less rough surface condition, which has the effect of improving the regermination of smaller columns and of intercolumnar spaces that are increasingly tighter as the upper layers are formed. These compressions result in: a limitation in the infiltration of the CMAS oxides; an increase in the service life of the thermal barrier system, and an increase in the tenacity of the thermal barrier system.

(21) In a third embodiment, shown in FIGS. 5 and 6, the bonding layer of the thermal barrier system of the two alternatives of the first embodiment is also subjected to a partial or total compression. Thus, the thermal barrier system has both the bonding layer thereof and the upper ceramic layer thereof placed in compression. FIG. 5 shows the thermal barrier system with a single ceramic layer 3, while FIG. 6 shows the thermal barrier system with n ceramic layers 31, . . . , 3i, . . . , 3n.

(22) The substrate 1 of the turbine blade is coated beforehand with a bonding layer 2 of the MCrAlY type, M designating Fe, Ni, Co and mixtures thereof. The bonding layer 2 can be obtained by conventional plasma spraying, for example of the APS (Air Plasma Spraying) type. The bonding layer 2 of the MCrAlY type can be replaced with a nickel aluminide or modified with platinum, or with a layer of the gamma/gamma-MCrAlY type.

(23) The compression of the bonding layer 2 makes it possible for: a partial or total densification of the bonding layer 2 and a control of the roughness thereof, with the benefit of the reduction in the deformation kinetics of this layer during the thermal cycle; the generation of residual stresses that have the effect of an increase in the hardness of the layer 2.

(24) In a fourth embodiment, shown in FIG. 7, the bonding layer 2 and the penultimate upper ceramic layer 3(n1) are partially or entirely subjected to a compression C2 and C3(n1) respectively.

(25) Finally, in a fifth embodiment, the bonding layer 2 and all of the ceramic layers 31, . . . , 3i, . . . , 3n are subjected to a compression (FIG. 8).