Turbine engine part coated with a protective ceramic coating, method for manufacturing and for using such a part

Abstract

A turbine engine part includes at least a substrate, and present on the substrate, a ceramic coating for protection against calcium and magnesium aluminosilicates, the ceramic coating including Al.sub.2O.sub.3 at a molar content lying in the range 33% to 49%, Y.sub.3Al.sub.5O.sub.12 at a molar content lying in the range 21% to 53%, and yttria-stabilized zirconia at a molar content lying in the range 13% to 31%.

Claims

1. A turbine engine part comprising at least a substrate, and present on the substrate, a ceramic coating for protection against calcium and magnesium aluminosilicates, the ceramic coating comprising: Al.sub.2O.sub.3 at a molar content lying in the range 33% to 49%; Y.sub.3Al.sub.5O.sub.12 at a molar content lying in the range 21% to 53%; and yttria-stabilized zirconia at a molar content lying in the range 13% to 31%.

2. A part according to claim 1, wherein the ceramic coating comprises: Al.sub.2O.sub.3 at a molar content lying in the range 37% to 45%; Y.sub.3Al.sub.5O.sub.12 at a molar content lying in the range 29% to 45%; and yttria-stabilized zirconia at a molar content lying in the range 17% to 27%.

3. A part according to claim 1, wherein the ceramic coating has a thickness lying in the range 50 m to 200 m.

4. A part according to claim 1, wherein the substrate comprises a material selected from the following: a metal superalloy and a ceramic matrix composite material.

5. A part according to claim 1, further comprising a thermal barrier layer present between the substrate and the ceramic coating.

6. A part according to claim 1, wherein the substrate constitutes a part for an aviation turbine engine selected from among the following parts: a turbine blade; at least a portion of a turbine nozzle; and at least a portion of a turbine ring.

7. A method of fabricating a part according to claim 1, the method comprising at least a step of forming the ceramic coating on the substrate, the ceramic coating comprising: Al.sub.2O.sub.3 at a molar content lying in the range 33% to 49%; Y.sub.3Al.sub.5O.sub.12 at a molar content lying in the range 21% to 53%; and yttria-stabilized zirconia at a molar content lying in the range 13% to 31%.

8. A method according to claim 7, wherein the ceramic coating is formed by sintering.

9. A method according to claim 7, wherein the ceramic coating is formed on the substrate by directed solidification.

10. A method of using a part according to claim 1, comprising a step of using the part at a temperature higher than 1000 C. in an oxidizing medium and in the presence of calcium and magnesium aluminosilicates.

Description

BRIEF DESCRIPTION OF THE DRAWING

(1) Other characteristics and advantages of the present invention appear from the following description made with reference to the accompanying drawing which show embodiments having no limiting character. In the figures:

(2) FIGS. 1 and 2 are enlarged section views of the surfaces of turbine engine parts in two embodiments of the invention.

DETAILED DESCRIPTION OF THE INVENTION

(3) FIG. 1 is an enlarged section view of the surface of a turbine engine part 10 comprising a substrate 11 coated in a ceramic coating 12 for providing protection against calcium and magnesium aluminosilicates (CMAS). The part 10 comprises in this order: the substrate 11, an adhesion layer 13, a thermal barrier layer 14, and the ceramic coating 12.

(4) In accordance with the invention, the ceramic coating 12 comprise alumina (Al.sub.2O.sub.3) having a molar content lying in the range 33% to 49%, and more particularly in the range 37% to 45%. The ceramic coating 12 also comprises yttrium aluminum garnet (Y.sub.3Al.sub.5O.sub.12 or YAG) and has a molar content lying in the range 21% to 53%, more preferably lying in the range 29% to 45%. The ceramic coating 12 also comprises yttria-stabilized zirconia (YSZ) at a molar content lying in the range 13% to 31%, and more preferably lying in the range 17% to 27%. The ceramic coating 12 may comprise only YAG, alumina, and YSZ. In other words, the ceramic coating 12 need not have any compound other than YAG, alumina, or YSZ.

(5) By way of example, the substrate 11 may comprise a metal superalloy, e.g. a nickel-based superalloy, or a ceramic matrix composite material.

(6) In the example shown in FIG. 1, the ceramic coating 12 is directly in contact with the thermal barrier layer 14, and is present on the surface of that layer 14. By way of example, the thickness e of the coating 12 may lie in the range 50 m to 200 m. The adhesion layer 13 is present between the substrate 11 and the thermal barrier layer 14, and is directly in contact therewith.

(7) In known manner, the thermal barrier layer 14 may comprise yttria stabilized zirconia (YSZ), which presents a columnar structure.

(8) The adhesion layer 13, which is itself known, serves to provide good adhesion between the thermal barrier layer 14 and the substrate 11. More generally, such an adhesion layer 13 serves to provide good mechanical compatibility between the thermal barrier layer 14 and the substrate 11, in particular serving to compensate for any differential thermal expansion that might exist between the materials of the layer 14 and of the substrate 11.

(9) When the substrate 11 comprises a metal superalloy, the adhesion layer 13 may for example comprise a simple or a modified aluminide (e.g. NiCrAlY for a nickel-based superalloy substrate), that may oxidize in part or that may form an oxide layer (also known as a thermally grown oxide (TGO) layer).

(10) When the substrate 11 comprises a ceramic matrix composite material, the adhesion layer 13 may comprise silicon.

(11) In general manner, the adhesion layer 13 is adapted as a function of the material making up the substrate 11 and the thermal barrier layer 14.

(12) FIG. 2 is an enlarged section view of the surface of a turbine engine part 20 in another embodiment of the invention. In this example, the part 20 comprises in this order: a substrate 21, an adhesion layer 23, and a ceramic coating 22. Unlike the example of FIG. 1, the part 20 in this example does not have an additional thermal barrier layer. Specifically, in this example, the ceramic coating 22 acts as a thermal barrier that protects the substrate 21.

(13) The substrate 21, the adhesion layer 23, and the ceramic coating 22 may present characteristics that are identical respectively to the substrate 11, to the adhesion layer 13, and to the ceramic coating 12 of the example of FIG. 1. Nevertheless, it is possible to modify the thickness e of the ceramic coating 22, e.g. by increasing it compared with the example of FIG. 1.

(14) In the example of FIG. 2, the adhesion layer 23 is present between the substrate 21 and the ceramic coating 22, and is directly in contact therewith. Like the part 10 of FIG. 1, the material forming the adhesion layer 23 may be adapted as a function of the material forming the substrate 21.

(15) In the example shown, the substrates 11, 21 of the part 10, 20 may consist in an aviation turbine engine part selected from the following: a turbine blade, at least a portion of a turbine nozzle, at least a portion of a turbine ring.

(16) Naturally, the invention is not limited to the configurations described above with reference to the examples shown, and it is entirely possible to envisage using other configurations of parts coated in a ceramic coating in the context of the invention.

(17) The present invention also provides a method of fabricating a part 10, 20 of the invention. Such a method comprises at least a step that consists in forming a ceramic coating 12, 22 of the kind described above on the substrate 11, 21 of the part. The ceramic coating 12 may be formed directly on a thermal barrier layer 14 present on the substrate 11, or in a variant directly on an adhesion layer 23 present on the substrate 21.

(18) The ceramic coating 12, 22 may be formed by sintering a powder composition comprising appropriate quantities of alumina, yttria, and zirconia. Under such circumstances, it should be observed that yttrium aluminum garnet (YAG) and yttria-stabilized zirconia (YSZ) are formed as a result of the sintering step. It is also possible to sinter directly a mixture of powders that comprise alumina, YAG, and YSZ.

(19) In a variant, the ceramic coating 12, 22 may be formed by a method of directed solidification using laser melting. For this purpose, powders of alumina, yttria, and zirconia may be placed on the substrate and then the deposited suspension layer can be subjected to laser melting in order to obtain the ceramic coating. As for sintering, under such circumstances, yttrium aluminum garnet and yttria-stabilized zirconia are formed as a result of the directed solidification step. It is also possible to make use directly of a mixture of powders comprising alumina, YAG, and YSZ.

(20) The present invention also provides a method of using a part 10, 20 of the invention. Such a method comprises a step of using the part 10, 20 at a temperature higher than 1000 C. in an oxidizing medium and in the presence of calcium and magnesium aluminosilicates. Such conditions of use correspond substantially to the environmental conditions in which the hot portions of an aviation turbine engine operate in a desert environment (these conditions existing in particular within the turbines of the engine). When a part 10, 20 of the invention is used under such conditions, the ceramic coating 12, 22 reacts with the calcium and magnesium aluminosilicates that come into contact therewith to form a protective layer that is impermeable to calcium and magnesium aluminosilicates. Thus the substrate 11, 21 of the part 10, 20 is protected by the coating 12, 22 of the invention.