CN107002214B

CN107002214B - Method for coating a turbine blade

Info

Publication number: CN107002214B
Application number: CN201580065728.5A
Authority: CN
Inventors: F·阿马德; C·门克
Original assignee: Siemens AG
Current assignee: Siemens Energy Global GmbH and Co KG
Priority date: 2014-12-04
Filing date: 2015-11-19
Publication date: 2019-12-27
Anticipated expiration: 2035-11-19
Also published as: DE102014224865A1; EP3198049B1; CN107002214A; WO2016087215A1; EP3198049A1

Abstract

The invention relates to a method (2) for coating a turbine blade (1), the turbine blade (1) comprising an airfoil (4) and at least one platform (2) arranged at the 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), comprising the following method steps: -applying at least a first layer (22) of a coating (24) on the platform (2) on the airfoil side, and-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 region (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 region (6).

Description

Method for coating a turbine blade

Technical Field

The 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 region and at least one raised region adjoining the contact region, and the airfoil is terminated at the contact region.

Background

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

In particular in the case of providing power compensation for networks with strongly fluctuating loads, in which load changes frequently occur in gas turbines, special technical requirements are imposed on the structure of the gas turbine. The efficiency of a gas turbine increases with increasing compression pressure and increasing combustion temperature. For efficient operation, it is also possible to produce positive effects using as light a material as possible for the rotor blades of the compressor and turbine stages. Likewise, in order to improve the efficiency of the stator blades and rotor blades of the compressor and turbine stages, they can be optimized in terms of their shape in terms of flow technology. The desire for special shaping for the blade mounting and also in the case of rotor blades based on materials which are as light as possible, is faced, inter alia, by the requirements for stability and heat resistance during operation of the gas turbine.

The blade mounting of the turbine stages is subjected to particular thermal and mechanical loads due to the high forces or high torques of the combusted, expanding hot gases. In order to optimize the mechanical stability, the individual blades of the turbine stage are therefore each produced as a single-crystal component, if possible. The required heat resistance, which single crystal materials usually do not provide, is then usually achieved by additional coatings. Wherein the coating can also be applied in multiple layers as desired.

A frequently used method for this purpose is to coat the regions of the blade where the highest thermal loads occur with a thermal barrier or thermal shielding ("TBC"), which is held on the blade by an adhesive layer ("bonding"). The adhesion layer is often formed by a superalloy, for example a metal chromium aluminum yttrium alloy having nickel and/or cobalt as base metal. The adhesive layer is sprayed onto the monocrystalline material of the blade, mostly in a defined thickness, in a multilayer manner. Thereby, on the one hand, the adhesion of the TBC to the blade is improved, and on the other hand the adhesion layer itself also contributes to an increased heat resistance of the blade. In the context 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 is paid to the specific structure of the turbine blade. The turbine blade is generally composed of a profiled airfoil which terminates at its two longitudinal ends in a platform. Wherein the airfoils are arranged in the radial direction in the flow space of the turbine, the platforms of the individual blades of the same turbine stage form an inner or outer ring and are correspondingly connected to one another as flush as possible. When spraying the airfoil sides of the platform with a layer of adhesive, the end faces of the platform are also often lightly sprayed. Such residues on the end faces of the platforms are undesirable because the distance between two adjacent platforms may be too large due to residual inhomogeneities, since these two platforms should be connected to one another as closely as possible to the tolerance in order to prevent flow losses as much as possible due to the gap between the two platforms.

For this reason, the remainder of the adhesive layer is 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 platform end face jeopardizes the coating on the airfoil side. Here, the adhesive layer may be broken or torn at the edge of the stage. If this occurs at the location where the TBC should be applied, it will deteriorate its adhesion. In operation, the TBC may gradually break off. If this occurs where the TBC is not located, the blade material is already subjected to higher thermal loads due to the cracks themselves.

Disclosure of Invention

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

The object is achieved 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 region and at least one planar protruding region adjoining the contact region, and the airfoil is terminated in the contact region. The method comprises the following method steps:

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

-removing at least the first layer of the coating from at least one end face of the platform in the region of the protruding areas, with the at least the first layer of the coating remaining on the end face in the region of the contact areas.

The airfoil can in particular be connected at both ends to in each case one platform, wherein in particular the method steps mentioned can be carried out on both platforms. The invention is based on the following considerations:

for fluid-technical reasons and for mechanical reasons, the two adjacent platforms of each two adjacent turbine blades are arranged as closely as possible to each other. If a coating is applied to the platform of the turbine blade on the airfoil side in this case, for example in order to improve the heat resistance, undesirable wetting or end coating of the platform can also occur depending on the particular process used for the coating. In this case, however, a controlled coating of the end face, that is to say an irregular coating, generally does not have a precisely defined thickness, does not generally occur. In this case, at certain positions of the two opposing end faces, irregularities of the coating may be superfluous, so that an undesired gap may form between the two adjacent platforms.

This can now be avoided by preventing the coating from being applied to the end faces. But this is very complicated according to the specific technical solution of the coating process. For process-related reasons, it is therefore not desirable to prevent at least partial application of the coating to the end faces at the same time as a corresponding effort. It is likewise not cost-intensive to completely remove the individual layers of the coating from the end face of the platform. In addition, damage to the layers of the platform on the respective airfoil side of the coating can occur here in particular on the edges.

It is now an important, entirely unexpected recognition of the present invention that the platform exhibits greater thermal expansion in the region of the protruding zone than in the region of the contact zone when the turbine is operating.

This has several reasons: on the one hand, the platform in the protruding region is subjected to the high temperatures occurring during operation over its entire surface expansion, whereas the platform is not subjected to these temperatures in the contact region at the location of the connecting airfoil. If, in the case of a predetermined temperature, it is assumed that all microscopic surface elements of the platform undergo a uniform thermal expansion, this means, in short, that a macroscopically greater thermal expansion occurs in the protruding region, since all microscopic surface elements are subjected to thermal effects, which is not the case for some microscopic surface elements in the contact region.

This effect is further enhanced if the microscopic area elements of the platform in the bulge region, immediately surrounding the region where the platform terminates the airfoil, expand less thermally at a set temperature than in other planar regions. At the location of the platform connection with the airfoil or transition to the airfoil, the turbine blade has a very high stability similar to a T-beam. This also affects the thermal expansion in this region.

In the region of the contact zone or the raised zone, the so-called differential thermal expansion of the platform is used by the invention in that at least a first layer of the coating is removed from the end face of the platform only in the region of the raised zone, while the corresponding layer is retained on the end face of the platform in the region of the contact zone. This is preferably carried out on all end faces of the blade installation of the turbine stage, each directly opposite the end face of the other platform.

In this case, during operation of the turbomachine, the platform expands 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 the adjacent platform is reduced. By leaving a layer of the coating on the end faces of the lands in the region of the contact zone, it is possible to avoid excessive clearance of adjacent lands in the region of the contact zone due to a stronger thermal expansion in the region of the protruding zones. The distance between two adjacent platforms can therefore be selected such that, during operation, they come into contact with their end faces in the region of the raised area quickly or completely, thereby preventing the leakage of hot gas by fluid technology. In the region of the contact region of two adjacent platforms, the leakage is impeded by a layer of the coating remaining on the end face.

In the context of the transition of the platform into the airfoil in the contact region by forming a concave surface, the retention of at least a first layer of the coating on the end face in the region of the contact region is particularly advantageous in a wide variety of possible geometries of the turbine blade. Removing portions of the coating may result in tangential forces within the layers of the coating. While such tangential forces can propagate largely unhindered in the convex region in the layer of the coating, they can locally have a normal component in the region of the contact region on the concave surface. This normal force component facilitates local separation of the layer of the coating concerned from the platform in the region of the concave surface and the contact zone. The risk of local separation of the layers can therefore be significantly reduced by dispensing with the removal of at least the first layer of the coating in the end face in the region of the contact region.

Preferably, at least a first layer of the coating is applied on the platform on the airfoil side by spraying. In this case, the method presented is particularly advantageous, since spraying the platform with the coating on the airfoil side can easily occur (possibly without control) wetting the end faces of the platform with a layer of the coating. Thus satisfying the preconditions of the method.

Alternatively, at least a first layer of the coating is applied to the platform by means of an impregnation bath. In particular in turbine blades in which the airfoil terminates at both ends in a respective platform, it is only difficult in this case to prevent the end faces of the platforms from being coated with a layer of coating. The preconditions for the method are therefore also met here.

Advantageously, a bonding layer is applied as a first layer of coating on the platform on the airfoil side. Such a bonding layer serves to optimize the bonding of further, later applied layers of the coating to the material of the turbine blade. The turbine blade can therefore be coated, in particular on the platform, with a material which can be optimized with respect to heat resistance and corrosion resistance. With this optimization, it is not necessary to consider the material separately for attachment to the raw material of the turbine blade, since this attachment is ensured by the adhesive layer.

The method is particularly advantageous if such a bonding layer forms the first layer of the coating on the platform on the airfoil side, since the bonding layer can be left intact in the sensitive region of the contact region. In this case, there is therefore no risk that the remaining adhesive layer on the end face may be damaged in the region of the contact region on the platform on the same airfoil side as a result of the removal, which would otherwise reduce the adhesion of further layers of the coating in this region.

Here, a superalloy is suitably applied as an adhesion layer on the platform. Such superalloys may be, inter alia, metal chromium aluminum yttrium compounds (MCrALY), in which nickel and/or cobalt may be used as base metals. Superalloys have very good properties with respect to their adhesion to the raw materials typically used for turbine blades.

It has further been found to be advantageous to remove at least the first layer of the coating from at least one end face of the platform by grinding in the region of the raised areas. This removal can be controlled particularly precisely locally, for example in comparison with an erosion method, as a result of which the risk of undesired damage to the layer of the coating can be reduced.

In a further embodiment of the invention, a heat shield is applied to the platform on the airfoil side as a further layer of the coating. This can be designed in particular as a ceramic thermal barrier. The use of the method is particularly advantageous in this case, since the risk of damaging the layers of the coating applied before the thermal barrier layer, in particular in the sensitive region of the contact region, can thereby be significantly reduced. This has a positive effect in particular on the adhesion of the thermal barrier.

The invention further relates to a gas turbine comprising at least one guide blade and/or rotor blade, which is coated by the aforementioned method. The advantages given for the method and its embodiments can be transferred to a gas turbine in analogy.

Drawings

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

fig. 1 shows an oblique view of the platform of a turbine blade, with an airfoil tip implicitly represented,

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

FIG. 3 shows a block diagram of a flow of a method for coating a turbine blade,

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

Mutually corresponding parts and dimensions are provided with the same reference numerals throughout the figures.

Detailed Description

Fig. 1 schematically shows an oblique view of the end of a turbine blade 1. Therein, a turbine blade 1 has a platform 2 and an airfoil 4, wherein the airfoil 4 is represented implicitly in this illustration as an airfoil tip. The region of the platform 2 that contacts the airfoil 4 or transitions into the airfoil 4 is defined herein as a contact region 6. This is indicated by the dotted border. At the contact region 6, on the platform 2, the airfoil 4 adjoins in both directions a projecting bulge region 8a, 8b, respectively. These projections 8a, 8b are each marked here by a dashed border. If the first layer of the coating is applied by spraying on the turbine blade 1, then it is generally not possible to avoid wetting the end face 10 of the platform 2 with a portion of the coating also when spraying the platform 2 on the airfoil side. In the region of the raised areas 8a, 8b, the remnants of the first layer of coating are removed from the end faces (10a, 10 b). In contrast, in the region of the contact region 6, the coating portion which reaches the end face 10c when the first layer is applied remains there.

Fig. 2 schematically shows a top view of two adjacent platforms 2 of a turbine blade 1. In this case, 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, a portion of which also falls on the respective end face 10 of the platform 2, this may result in the gap 12 separating the two platforms 2 from one another no longer having a defined width. In order to prevent this, the residues of the layer of applied coating on the end faces 10a, 10b are removed in the region of the raised areas 8a, 8b of the two platforms 2. In the region of each platform 2, which is connected to the contact region 6 of the airfoil 4 on the platform 2, the coating remains on the end face 10 c.

During operation of the gas turbine with the turbine blade 1, the platform 2 expands more strongly thermally in the protruding regions 8a, 8b than in the contact region 6. This results in a gap 12 between two adjacent platforms 2 in the region of the projecting zones 8a, 8b having a smaller spacing during operation. An excessive width of the gap 12 in the region of the contact zone 6, which could lead to a fluidic undesirable leakage of hot gas, can therefore be avoided by the remaining layer of coating on the end face 10c in this region.

A block diagram of a 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. The platform 2 of the turbine blade 1 is also partially coated on the end face 10 of the platform 2 by spraying 26 it with a superalloy 28.

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 raised regions 8 a. In the region of the bulging region 8a, the platform is coated with the superalloy 28 after grinding 30 only on the airfoil side and not on the end face 10 a. In the region of the contact region, which is not shown in detail, the superalloy 28 remains there as well on the end face.

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

Fig. 4 schematically shows a cross-sectional illustration of a gas turbine 40 with turbine blades 1, which turbine blades 1 are coated according to the method described above. The turbine blades 1 can be configured here not only as guide blades 42 but also as rotor blades 44.

While the invention has been further illustrated and described in detail by the preferred embodiments, it is not to be restricted by the embodiments. Other variants can be derived therefrom by the skilled person without departing from the scope of protection of the invention.

Claims

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.