CN107002214B - Method for coating a turbine blade - Google Patents

Abstract

The invention proposes a method (2) for coating a turbine blade (1) comprising an airfoil (4) and at least one platform arranged at the end of the airfoil (4) (2), wherein the or each platform (2) has a contact area (6) and at least one planar projection (8a, 8b) adjoining the contact area (6), and makes the airfoil ( 4) Terminated in this contact zone (6), with the following method steps: applying at least a first layer (22) of the coating (24) on the platform (2) on the airfoil side, and applying the coating (24) In the case where the at least first layer (22) of the coating (24) remains on the end face (10c) within the range of the contact area (6), the at least first layer (22) of the coating (24) is in the raised area (8a, 8b) is removed from at least one end face (10a, 10b) of the platform (2).

Description

Method for coating turbine blades

technical field

The present invention relates to a method for coating a turbine blade comprising an airfoil and at least one platform arranged at the end of the airfoil, wherein the or each platform has a contact area and at least one contact area with adjoining the bulge area, and the airfoil terminates in this contact area.

Background technique

Step-by-step conversion of energy production to increased renewable energy carriers presents many technical challenges. Due to the very variable availability, especially in many places of wind and solar, which represent the two most important renewable energy carriers, there is a need to equalize fluctuations in the input power for stable network operation with constant supply power . In this context, gas turbines play a key role due to their high flexibility compared to other conventional energy carriers.

Especially in the case of power compensation for networks with strongly fluctuating loads, where load changes frequently occur in the gas turbine, special technical requirements are placed on the structure of the gas turbine. The efficiency of a gas turbine increases with increasing compression pressure and with increasing combustion temperature. For efficient operation, the use of the lightest possible material for the rotor blades of the compressor and turbine stages can also have a positive effect. Likewise, in order to improve the efficiency of the guide vanes and rotor blades of the compressor and turbine stages, it is possible to optimize their shape in terms of fluid technology. Among other things, the desire for a special profile for blade mounting, and also in the case of rotor blades from the lightest possible material, is faced with the requirements for stability and thermal resistance during operation of the gas turbine.

Due to the high force effects or high torques of the combusted, expanding hot gases, the blade installation of the turbine stage is subjected to particular thermal and mechanical loads. Therefore, in order to optimize the mechanical stability, the individual blades of the turbine stage are manufactured individually as single crystal components, if possible. The required thermal resistance, which is usually not provided by single crystal materials, is then usually achieved by additional coatings. Here, the coating can also be applied in multiple layers as required.

The method often used for this is to coat the areas of the blade where the highest thermal load occurs with a thermal barrier or thermal shield (“therma barrier coating”, TBC), which is passed through an adhesive layer ("therma barrier coating", TBC). "bondcoating") remains on the blade. In this case, the adhesive layer is often formed by superalloys, for example metallic chromium-aluminum-yttrium alloys with nickel and/or cobalt as base metals. Among them, the adhesive layer is mostly sprayed on the single crystal material of the blade in a multi-layered form with a defined thickness. As a result, on the one hand the adhesion of the TBC to the blade is improved, and on the other hand the adhesive layer itself likewise contributes to an increased heat resistance of the blade. In the range of larger-scale thermal loads, the adhesive layer alone can also exhibit sufficient thermal protection for the material of the turbine blade.

In the coating, attention needs to be paid to the special structure of the turbine blades. The turbine blade typically consists of a profiled airfoil that terminates in a platform at each of its two longitudinal ends. In this case, the airfoils are arranged radially in the flow space of the turbine, the platforms of the individual blades of one and the same turbine stage form an inner or outer ring and are respectively connected to one another as flush as possible. When spraying the airfoil sides of the platform with the adhesive layer, the end faces of the platform are also often lightly sprayed. Since two adjacent platforms should be connected to each other as closely as possible to tight tolerances in order to prevent as much as possible flow losses caused by the gap between the two platforms, such residues on the end faces of the platforms are undesirable because due to residual The non-uniformity of , may lead to excessive spacing between these two adjacent platforms.

For this reason, the remnants of the adhesive layer are removed from the end face of the platform, which is usually done manually, for example by grinding. This is complicated on the one hand, and on the other hand, the grinding process on the end face of the platform jeopardizes the coating of the airfoil side. Here, the adhesive layer may break or tear at the edges of the platform. If this occurs where the TBC should be coated, it worsens its adhesion. During operation, the TBC may gradually come off. If this occurs where no TBC is provided, the blade material is already subject to higher thermal loads due to the cracks themselves.

SUMMARY OF THE INVENTION

The object of the present invention is therefore to propose a method for coating turbine blades which is as easy to carry out as possible and achieves the best possible thermal and corrosion protection in the region of the platform.

Said task is solved according to the invention by a method for coating a turbine blade comprising an airfoil and at least one platform arranged at the end of the airfoil, wherein the or each platform has a contact area and At least one planar bulge region adjoins the contact region and terminates the airfoil in the contact region. Here, the method has the following method steps:

– applying at least a first layer of coating to the platform on the airfoil side, and

- Removing at least the first layer of the coating from the at least one end face of the platform in the region of the protruding zone, while retaining the at least first layer of the coating on the end face in the region of the contact zone.

In particular, the airfoil can be connected at each end to a platform, in which case the mentioned method steps can be carried out in particular on both platforms. The present invention is based on the following considerations:

For fluid-technical reasons as well as for mechanical reasons, two adjacent platforms of every two adjacent turbine blades are arranged as closely as possible to each other. If a coating is then applied to the platform of the turbine blade, on the airfoil side, for example to improve thermal resistance, undesired wetting or coating of the end face of the platform may also occur, depending on the specific process used for the coating. Floor. In this case, however, a controlled coating of the end face usually does not occur, rather the coating is irregular and therefore usually does not have a precisely defined thickness. Here, at certain positions of the two opposite end faces, irregularities of the excess coating may be involved, so that an undesired gap may be formed between the two adjacent platforms.

This can now be avoided by preventing the application of a coating on the end faces. But this is very complicated depending on the specific technical solution of the coating process. Therefore, for technical reasons, it is not desired to prevent at least partial application of the coating on the end face with relative effort. The complete removal of the individual layers of the coating from the end face of the platform is equally costly. Furthermore, damage to the individual airfoil-side layers of the coating of the platform can occur here, in particular at the edges.

An important and completely unexpected realization of the present invention is now that the platform exhibits a greater thermal expansion in the region of the bulge than in the region of the contact region during operation of the turbine.

There are several reasons for this: on the one hand, in the bulge region, the platform is exposed to the high temperatures that occur during operation in the expansion of its entire area, whereas the platform is not exposed to this temperature in the contact region at the point where the airfoil is attached. If, at a preset temperature, it is assumed that all microscopic area elements of the platform exhibit uniform thermal expansion, then in simple terms, this means that macroscopic thermal expansion occurs in the bulge, because here all Microscopic area elements suffer from thermal effects, which is not the case for some microscopic area elements within the contact region.

This effect is further enhanced if the thermal expansion of the microscopic area elements of the platform in the bulge region, immediately around the region where the platform terminates the airfoil, is less at the set temperature than in other planar regions . Where the platform joins or transitions to the airfoil, the turbine blade has a very high stability similar to a T-beam. This also affects thermal expansion in this area.

In the area of the contact area or the projection area, the so-called differential thermal expansion of the platform is utilized here by the invention by removing at least the first layer of the coating from the end face of the platform only in the area of the projection area, while the The corresponding layer is retained on the end face of the platform in the region of the contact zone. Therein, this is preferably carried out on all end faces of the turbine stage which are each directly opposite the end face of the other platform in the blade installation.

In this case, during operation of the turbine, the platforms expand more strongly in the region of the protruding region than in the region of the contact region, as a result of which the distance to adjacent platforms is reduced. By retaining a layer of coating on the end face of the platform in the region of the contact region, it is possible to avoid the formation of an oversize of the adjacent platform in the region to the contact region due to a stronger thermal expansion in the region of the projection region Clearance. Therefore, the distance between two adjacent platforms can be selected such that during operation, their end faces in the region of the bulges are contacted quickly or completely, thereby preventing the leakage of hot gas by fluid technology. In the region of the contact area of two adjacent platforms, this leakage is prevented by the layer of the coating which remains on the end face accordingly.

In the context of the transition of the platform to the airfoil in the contact zone by forming a concave surface under the various possible geometries of the turbine blade, it is particularly advantageous to retain at least a first layer of coating on the end face in the region of the contact zone. Removing part of the coating may result in tangential forces within the layers of the coating. While this tangential force can propagate largely unhindered in the protruding region in the layers of the coating, it can locally have a normal direction on the concave surface in the region of the contact region weight. This normal force component facilitates the local separation of the layers of the coating concerned from the platform in the region of the concave surface and the contact zone. Thus, the risk of local separation of the layers can be significantly reduced by not removing at least the first layer of the coating in the end face in the region of the contact zone.

Preferably, at least a first layer of the coating is applied on the platform by spraying on the airfoil side. In this case, the presented method is particularly advantageous, since spray coating the platform on the airfoil side can easily (possibly without control) wet the end face of the platform with a layer of coating. The prerequisites for this method are therefore met.

Alternatively, at least a first layer of the coating is applied to the platform by means of an impregnation bath. Especially in turbine blades in which the airfoil terminates at both ends in the respective platform, it is only in this case difficult to prevent the application of a coating layer on the end face of the platform. Therefore, the preconditions for the method are also fulfilled here.

Advantageously, an adhesive layer is applied on the platform as the first layer of the coating on the airfoil side. This bonding layer is used to optimize the bonding of layers of other, later applied coatings to the raw material of the turbine blade. Thus, the turbine blade, especially on the platform, can be coated with a material which can be optimized with regard to thermal and corrosion resistance. With this optimization, the attachment of the material to the raw material of the turbine blade does not have to be considered separately, since the attachment is ensured by the adhesive layer.

The method is particularly advantageous if such an adhesive layer forms the first layer of the coating on the platform on the airfoil side, since the adhesive layer can be left intact in the sensitive area of the contact area. Therefore, there is no risk that the end face could be damaged due to the removal of the remaining adhesive layer on the same airfoil side in the region of the contact zone on the platform, which would otherwise lead to a reduction of the coating in this region the attachment of other layers.

Here, the platform is expediently coated with a superalloy as an adhesive layer. Such superalloys can in particular be metallic chromium aluminium yttrium compounds (MCrALY), wherein nickel and/or cobalt can be used as base metals. Superalloys have very good properties with regard to their adhesion to the raw materials typically used for turbine blades.

It has further been shown that it is advantageous to remove at least the first layer of the coating from the at least one end face of the platform by grinding in the region of the raised area. This type of removal, for example, can be controlled particularly locally in comparison with erosion methods, whereby the risk of undesired damage to the layers of the coating can be reduced.

In another technical solution of the present invention, a heat shield layer is coated on the platform on the airfoil side as another layer of the coating. In particular, this can be designed as a ceramic heat shield. The application of the method is particularly advantageous in this case, since the risk of damaging the layers of the coating applied before the heat shield layer can thereby be significantly reduced, especially in sensitive areas of the contact zone. This has a positive effect in particular on the adhesion of the heat shield.

Furthermore, the invention relates to a gas turbine comprising at least one guide vane and/or a rotor blade, which is coated by the aforementioned method. Among them, the advantages given for the method and its extension can be transferred to gas turbines by analogy.

Description of drawings

Embodiments of the present invention are further described below with reference to the accompanying drawings. in:

Figure 1 shows an oblique view of the platform of a turbine blade, with the airfoil tips implied,

Figure 2 shows a top view of two adjacent platforms of a turbine blade,

Figure 3 shows a block diagram of the flow of a method for coating a turbine blade,

Figure 4 shows a schematic cross-sectional view of a gas turbine.

Corresponding parts and dimensions are assigned the same reference numerals in all figures.

Detailed ways

An oblique view of the end of a turbine blade 1 is schematically shown in FIG. 1 . Therein, the turbine blade 1 has a platform 2 and an airfoil 4, wherein the airfoil 4 is implicitly indicated as an airfoil tip in the schematic diagram. The area of the platform 2 that is in contact with or transitions to the airfoil 4 is defined here as the contact zone 6 . This is indicated by a dotted border. At the contact zone 6, on the platform 2, the airfoils 4 adjoin the protruding protruding zones 8a, 8b, respectively, in both directions. These raised areas 8a, 8b are each marked here by dashed borders. If the first layer of the coating is then applied by spraying on the turbine blade 1 , it is generally unavoidable to also wet the end face 10 of the platform 2 with a portion of the coating when the platform 2 is sprayed on the airfoil side. In the region of the raised areas 8a, 8b, residues of the first layer of the coating are removed from the end faces (10a, 10b). In contrast, in the region of the contact zone 6 , the portion of the coating that reaches the end face 10c during the application of the first layer remains there.

A top view of two adjacent platforms 2 of a turbine blade 1 is schematically shown in FIG. 2 . Here, each of the two platforms 2 has an end face 10 which is opposite the end face 10 of the respective other platform. If a layer of coating is applied to the turbine blade 1 , wherein part of the layer of coating also falls on the corresponding end face 10 of the platform 2 , this may result in the gap 12 separating the two platforms 2 from being incompatible with each other. Again have a defined width. In order to prevent this, in the region of the raised areas 8a, 8b of the two platforms 2, residues of the layers of the applied coating on the end faces 10a, 10b are removed. In the region of each platform 2 , corresponding to the contact region 6 of the airfoil 4 on the platform 2 , residues of the layer of coating respectively remain on the end face 10 c .

During operation of the gas turbine with the turbine blades 1 , the platform 2 is thermally expanded more strongly in the protruding regions 8 a , 8 b than in the contact region 6 . This results in a smaller spacing of the gaps 12 between two adjacent platforms 2 in the region of the projections 8a, 8b during operation. As a result of the remaining layer of coating on the end face 10c in the region of the contact zone 6, it is possible to prevent the gap 12 from having an excessively large width in this region, which could lead to undesired leakage of hot gases in fluid technology.

A block diagram of the flow of a method 20 for coating a turbine blade 1 is schematically shown in FIG. 3 . On the turbine blade 1 , a first layer 22 of a coating 24 is first applied by spraying 26 on the airfoil side. The first layer 22 of the coating 24 is a superalloy 28, such as MCrALY. By spraying 26 the platform 2 of the turbine blade 1 with superalloy 28 it is also partially coated on the end face 10 of the platform 2 .

In a subsequent method step, the first layer 22 of the coating 24 applied to the end face 10a is removed from the end face 10a by grinding 30 in the region of the projection 8a. In the region of the bulge 8a, the platform after grinding 30 is coated with superalloy 28 only on the airfoil side and not on the end face 10a. In the region of the contact zone, which is not shown in detail, the superalloy 28 still remains there on the end face.

In another method step, another layer of coating 24 is applied on the airfoil side. The other layer is formed of ceramic TBC 32. For the TBC 32 , the superalloy 28 is intended to be the adhesive layer 34 , ie, the adhesion of the TBC 32 to the blade 1 is significantly improved by the superalloy 28 . In this context, it is particularly advantageous to leave the end face 10 free in the region of the contact region during the grinding 30 so as not to damage the first layer 22 of the coating 24 formed from the superalloy 28 in this sensitive region.

A schematic cross-sectional view of a gas turbine 40 with turbine blades 1 , which are coated according to the aforementioned method, is shown schematically in FIG. 4 . In this case, the turbine blades 1 can be configured not only as guide blades 42 but also as rotor blades 44 .

Although the present invention is further illustrated and described in detail by means of a preferred embodiment, the present invention is not limited by this embodiment. Other variants can be derived therefrom by the skilled person without departing from the scope of protection of the present invention.

Claims (8)

1. A method (20) for coating a turbine blade (1), the turbine blade (1) comprising an airfoil (4) and at least one platform (2) arranged at an end of the airfoil (4), wherein the or each platform (2) has a contact region (6) and at least one planar protruding region (8a, 8b) adjoining the contact region (6), and the airfoil (4) is terminated at the contact region (6), having the following method steps:

-applying at least a first layer (22) of a coating (24) on the platform (2) on the airfoil side,

-removing the at least first layer (22) of the coating (24) from at least one end face (10a, 10b) of the platform (2) in the region of the protruding areas (8a, 8b) with the at least first layer (22) of the coating (24) remaining on the end face (10c) in the region of the contact area (6).

2. The method (20) of claim 1, wherein the at least a first layer (22) of the coating (24) is applied by spraying (26) on the airfoil side.

3. The method (20) of claim 1, wherein the at least a first layer (22) of the coating (24) is applied by an infiltration bath.

4. A method (20) according to any of claims 1 to 3, wherein a bonding layer (34) is applied on the platform (2) on the airfoil side as the first layer (22) of the coating (24).

5. The method (20) according to claim 4, wherein a superalloy (28) is coated on the platform (2) as an adhesion layer (34).

6. A method (20) according to any one of claims 1 to 3, wherein the at least first layer (22) of the coating (24) is removed from the at least one end face (10a, 10b) of the platform (2) by grinding in the region of the protruding areas (8a, 8 b).

7. A method (20) according to any of claims 1 to 3, wherein a heat shield layer (32) is applied on the platform (2) on the airfoil side as a further layer of the coating (24).

8. Gas turbine (40) comprising at least one guide blade (42) and/or rotor blade (44), the at least one guide blade (42) and/or rotor blade (44) being coated by a method according to any one of claims 1 to 7.