Thermal barrier coating of gas turbine

Abstract

The utility model discloses a thermal barrier coating of a gas turbine, which comprises a bonding layer and a ceramic surface layer, and the ceramic surface layer is attached to an alloy matrix of a turbine blade through the bonding layer. Grooves are formed in the ceramic surface layer, are distributed in a grid shape, penetrate through the ceramic surface layer and are cut into the bonding layer. The grooves are formed in the thermal barrier coating, so that a telescopic space is provided for thermal expansion of a ceramic surface layer material, the overall expansibility of the thermal barrier coating is better, namely, the overall thermal expansion coefficient of the thermal barrier coating is increased and is closer to that of an alloy base body of the turbine blade, and the thermal barrier coating is better in thermal expansion performance. Therefore, the difference of thermal expansion performance between the turbine blade and the thermal barrier coating is reduced, and the risk of cracking and falling of the coating is reduced.

Description

TECHNICAL FIELD

[0001] The utility model relates to gas turbine technical field, especially a gas turbine thermal barrier coating. BACKGROUND

[0002] With the great progress of our country's science and technology and the increasing pressure of energy saving and emission reduction, the gas turbine is widely used in the market due to its high efficiency and environmental friendliness. The turbine blade is a key hot component of the gas turbine, which plays a decisive role in the output power of the entire gas turbine. With the gradual increase of the turbine inlet temperature, which can now reach 1400 DEG C, the protection of the turbine blade under high temperature is particularly important, especially the increasing application of thermal barrier coating.

[0003] The gas turbine thermal barrier coating is a coating with certain heat insulation capacity prepared on the surface of the high-temperature alloy turbine blade, i.e. the thermal barrier coating. The thermal barrier coating is mostly made of ceramic material, which has low thermal conductivity, high melting point and good chemical stability, and can effectively insulate heat and resist high temperature.

[0004] However, the most significant problem of this thermal barrier coating is that the surface ceramic layer and the metal matrix of the turbine blade itself have very different physical properties. The biggest problem in the actual application of the gas turbine is the difference in thermal expansion coefficient. Since the ceramic coating is a non-metallic compound, the thermal expansion coefficient is much lower than that of the high-temperature alloy. In addition, the uneven heating during actual use further exacerbates the unevenness of the thermal expansion between the thermal barrier coating and the turbine blade, thus easily leading to the cracking and falling off of the coating, thereby losing the protective effect on the turbine blade. SUMMARY

[0005] The utility model discloses a gas turbine thermal barrier coating to solve the problems in the prior art.

[0006] The utility model discloses a gas turbine thermal barrier coating to solve the problems in the prior art.

[0007] Preferably, the material of the adhesive layer is aluminum nickel cobalt alloy, and the material of the ceramic surface layer is aluminum nickel cobalt alloy, aluminum oxide, lanthanum zirconium oxide or yttrium aluminum garnet.

[0008] Preferably, the depth of the groove cutting into the adhesive layer is not more than 1 / 3 of the thickness of the adhesive layer.

[0009] Preferably, the thickness of the ceramic surface layer is 2-2.5 times the thickness of the adhesive layer.

[0010] Preferably, the thickness of the adhesive layer is 100 microns, the thickness of the ceramic surface layer is 200 microns, and the depth of the groove cutting into the adhesive layer is 30 microns.

[0011] Preferably, the number of grooves on a section line of the turbine blade satisfies the following formula:

[0012] n = (aL1-aL2) * DeltaT * L / c

[0013] Wherein, n is the number of grooves on a section line of the turbine blade; aL1 is the linear expansion coefficient of the turbine blade base alloy; aL2 is the linear expansion coefficient of the ceramic surface layer; DeltaT is the temperature change value of the turbine blade; L is the length of a section line of the turbine blade; and c is the transverse width of the groove.

[0014] The turbine blade of the present application has the following advantages: the grooves provided on the thermal barrier coating provide expansion space for the thermal expansion of the ceramic surface layer, so that the overall expansion of the thermal barrier coating is better, i.e., the overall thermal expansion coefficient of the thermal barrier coating is improved, so that the overall thermal expansion coefficient of the thermal barrier coating is closer to the thermal expansion coefficient of the turbine blade alloy base, thereby reducing the difference in thermal expansion performance between the turbine blade and the thermal barrier coating, and further reducing the risk of coating cracking and falling off. BRIEF DESCRIPTION OF DRAWINGS

[0015] Fig. 1 The figure is a structural schematic diagram of the present application.

[0016] Fig. 2 The figure is a sectional view of the thermal barrier coating.

[0017] In the figure: 1, alloy base; 2, ceramic surface layer; 3, groove; and 4, adhesive layer. DETAILED DESCRIPTION

[0018] The technical solutions in the embodiments of the present application will be described clearly and completely below with reference to the drawings in the embodiments of the present application. Obviously, the described embodiments are only a part of the embodiments of the present application, rather than all the embodiments. Based on the embodiments in the present application, all other embodiments obtained by those skilled in the art belong to the scope of protection of the present application.

[0019] As shown in the figure, a thermal barrier coating of a gas turbine includes an adhesive layer and a ceramic surface layer, and the ceramic surface layer is attached to an alloy base of a turbine blade through the adhesive layer. Figs. 1-2 The ceramic surface layer is provided with grooves, and the grooves are distributed in a grid shape and penetrate the ceramic surface layer and cut into the adhesive layer.

[0020] The grooves are straight-line grooves, and the grooves are arranged in a transverse and longitudinal staggered manner to form a grid structure.

[0021] The utility model discloses a supersonic flame spraying equipment is used on alloy base body of turbine blade and sprays adhesive layer, and plasma spraying equipment is used on adhesive layer and sprays ceramic surface layer, and the purpose of adhesive layer is to improve the adhesion strength of ceramic surface layer, after the preparation of ceramic surface layer, the grid-shaped groove is engraved on ceramic surface layer through laser engraving process, and the grid-shaped groove completely penetrates ceramic surface layer and cuts into adhesive layer, and the grid-shaped groove covers the surface of turbine blade. By setting the groove on the thermal barrier coating, the expansion space is provided for the thermal expansion of the ceramic surface layer material, so that the overall expansion of the thermal barrier coating is better, that is, the overall thermal expansion coefficient of the thermal barrier coating is improved, so that the overall thermal expansion coefficient of the thermal barrier coating is more close to the thermal expansion coefficient of the turbine blade alloy base body, thereby reducing the difference between the thermal expansion performance of the turbine blade and the thermal barrier coating, and further reducing the risk of coating cracking and falling off.

[0022] Wherein, by adjusting the depth of the groove, the distribution density of the groove, the width of the groove can change the overall thermal expansion coefficient of the thermal barrier coating. In the actual process, by reasonably setting the depth of the groove, the distribution density of the groove, the width of the groove, the overall thermal expansion coefficient of the thermal barrier coating can be minimized.

[0023] Specifically, the number of grooves on a certain cross-sectional line of the turbine blade satisfies the following formula:

[0024] n=(alphaL1-alphaL2)*DeltaT*L / c;

[0025] Wherein, n is the number of grooves on a certain cross-sectional line of the turbine blade; alphaL1 is the linear expansion coefficient of the turbine blade base alloy, unit is 1 / DEG C; alphaL2 is the linear expansion coefficient of the ceramic surface layer, unit is 1 / DEG C; DeltaT is the temperature change value of the turbine blade, unit is DEG C; L is the length of a certain cross-sectional line of the turbine blade, unit is mm; C is the transverse width of the groove, unit is mm.

[0026] If L in the above formula is the axial length of the turbine blade, the number of grooves in the axial direction of the turbine blade (i.e. in the direction from the root to the top of the blade) can be obtained.

[0027] In the utility model, the laser engraving process is used to process the groove, and the processing technology belongs to non-contact processing, which has the advantages of no processing stress and small size limit, can process very small grooves, and can process micron-level grooves.

[0028] In the embodiment, the material of the adhesive layer is aluminumized nickel-cobalt alloy, and the material of the ceramic surface layer is aluminumized nickel-cobalt alloy or aluminum oxide or lanthanum-zirconium oxide or yttrium-aluminum garnet.

[0029] The depth of the groove cutting into the bonding layer is not more than 1 / 3 of the thickness of the bonding layer. The depth of the groove is greater than the thickness of the ceramic face layer and is less than or equal to the thickness of the entire thermal barrier coating (bonding layer + ceramic face layer). If the depth of the groove is too shallow, the fragmentation of the ceramic face layer cannot be achieved, which is not conducive to the increase of the thermal expansion coefficient of the thermal barrier coating. If the depth of the groove is too deep, the protection of the turbine blade alloy substrate is not conducive.

[0030] The thickness of the ceramic face layer is 2-2.5 times the thickness of the bonding layer. In the embodiment, the thickness of the bonding layer is 100 μm, the thickness of the ceramic face layer is 200 μm, and the depth of the groove cutting into the bonding layer is 30 μm.

[0031] The utility model is not limited to the above-mentioned best implementation, and anyone can draw other various forms of products under the enlightenment of the utility model, but no matter make any change in its shape or structure, all have the same or similar technical scheme with the application, fall in the protection scope of the utility model.

Claims

1. A gas turbine thermal barrier coating, characterized by, The ceramic surface layer is attached to the alloy base of the turbine blade through the adhesive layer; the ceramic surface layer is provided with grooves, the grooves are distributed in a grid shape, and the grooves penetrate the ceramic surface layer and cut into the adhesive layer.

2. The thermal barrier coating for a gas turbine engine of claim 1, wherein, The material of the adhesive layer is an aluminized nickel-cobalt alloy, and the material of the ceramic surface layer is an aluminized nickel-cobalt alloy, alumina, lanthanum-zirconium oxide or yttrium-aluminum garnet.

3. The thermal barrier coating for a gas turbine engine of claim 1 wherein, The depth of the grooves cutting into the adhesive layer is not more than 1 / 3 of the thickness of the adhesive layer.

4. The thermal barrier coating for a gas turbine engine of claim 1 wherein, The thickness of the ceramic surface layer is 2-2.5 times the thickness of the adhesive layer.

5. The thermal barrier coating for a gas turbine engine of claim 4, wherein, The thickness of the adhesive layer is 100 microns, the thickness of the ceramic surface layer is 200 microns, and the depth of the grooves cutting into the adhesive layer is 30 microns.

6. The thermal barrier coating for a gas turbine engine of claim 1 wherein, The number of grooves on a cross-sectional line of the turbine blade satisfies the following formula: n=(alphaL1-alphaL2)*deltaT*L / c; wherein n is the number of grooves on a cross-sectional line of the turbine blade; alphaL1 is the linear expansion coefficient of the turbine blade base alloy; alphaL2 is the linear expansion coefficient of the ceramic surface layer; deltaT is the temperature change value of the turbine blade; L is the length of a cross-sectional line of the turbine blade; and c is the transverse width of the grooves.

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