EP3198049B1 - Method for coating a turbine blade - Google Patents

Description

The invention relates to a method of coating a turbine blade comprising a blade and at least one platform disposed at one end of the blade, the or each platform having a contact zone and at least one protrusion zone adjacent to the contact zone, and terminating the blade at the contact zone.

The gradual conversion of energy generation to increasingly renewable energy sources brings with it a multitude of technical challenges. As a result of the highly variable availability of, in particular, wind and solar energy, which are the two most important renewable energy sources in many places, the need for balancing the fluctuations in the power fed in is created for stable grid operation with a constant power supply. In this context, gas turbines have a key role to play because of their high flexibility compared to other conventional energy sources.

Not least for the provision of power balancing in networks with highly fluctuating load, in which take place in a gas turbine frequent load changes, so arise special technical requirements for the construction of the gas turbine. The efficiency of a gas turbine increases with increasing compression pressure and increasing combustion temperature. For efficient operation, the use of the lightest possible materials for the rotor blades in the compressor and turbine stages can also have a positive effect. Likewise, in order to improve the efficiency, the guide vanes and rotor blades of the compressor and turbine stages can be fluidly optimized in their shape. The desire for a special shaping for the blading, and in the case of the blades also for the lightest possible materials, are the requirements of the Strength and heat resistance in the operation of the gas turbine compared.

Due to the high force or the high torque due to the burned, expanding hot gas, the blading of a turbine stage is exposed to particular thermal and mechanical loads. The individual blades of a turbine stage are therefore made as a single-crystalline workpieces, if possible, to improve the mechanical strength. The required heat resistance, which is often unable to provide the monocrystalline material, is then usually achieved via an additional coating. Depending on the requirements, the coating can also be applied in several layers.

A frequently used method for this purpose is to coat the areas of a blade, in which the highest thermal loads occur, with a thermal barrier layer ("TBC"), which comprises a bond coating. is held on the blade. In this case, the bonding layer is usually formed by a superalloy, for example a metal-chromium-aluminum-yttrium alloy (MCrAlY) with nickel and / or cobalt as base metal. The bonding layer is usually sprayed onto the monocrystalline material of the blade with a defined thickness in several layers. In this way, on the one hand, the adhesion of the TBC to the blade is to be improved, on the other hand, the binding layer itself also contributes to the increased heat resistance of the blade. In areas of moderate thermal stress, the bond layer alone can provide sufficient heat protection for the turbine blade material.

When coating the special design of a turbine blade is to be observed. This usually consists of a profiled wing, which is completed at its two longitudinal ends in each case by a platform. The wing is arranged radially in the flow space of the turbine, the platforms individual blades of the same turbine stage form an inner or outer ring and close as flush as possible to each other. When spraying the respective wing-side surface of the platform with the bonding layer, there is often also a slight spraying of the respective end face of the platform. Since two adjacent platforms should join each other as closely tolerated as possible in order to prevent flow losses through the gap between the two platforms as possible, such remains at the end faces of the platforms are undesirable because they can lead to greater distances between the two adjacent platforms due to their irregularity.

For this reason, as the publication EP 2 366 488 A1 indicates that the remnants of the bonding layer away from the faces of the platforms, which is usually done manually, for example by grinding. On the one hand, this is expensive; on the other hand, the grinding process on the end face of a platform endangers the wing-side coating. The binding layer may break or tear at the edge of the platform. If this occurs at a point where a TBC is to be applied, its liability is impaired. During operation, the TBC may gradually flake off. If this occurs at a location where no TBC is provided, the material of the blade is already subject to a higher heat load by the cracks itself.

The invention is therefore based on the object of specifying a method for coating a turbine blade, which is as simple as possible to carry out, and in the area of the platform provides the best possible protection against heat and corrosion.

• Flügelseitiges Auftragen wenigstens einer ersten Lage einer Beschichtung auf die flächige Überstandszone und die Kontaktzone der Plattform und
• Entfernen der wenigstens ersten Lage der Beschichtung von wenigstens einer Stirnfläche der Plattform im Bereich der Überstandszone unter Belassen der wenigstens ersten Lage der Beschichtung an der gleichen Stirnfläche im Bereich der Kontaktzone.

Said object is achieved according to the invention by a method for coating a turbine blade, which comprises a wing and at least one arranged at one end of the wing platform, wherein the or each platform a contact zone and at least one of the contact zone has adjacent flat supernatant zone, wherein the supernatant zone in its entire surface extent the high temperatures occurring during operation can be exposed, and the contact zone, however, at the locations where the wing is connected to the platform and thus merges into these, these temperatures is not suspendable , The method has the following method steps:

• Wing-side application of at least a first layer of a coating on the planar overhang zone and the contact zone of the platform and
• Removing the at least first layer of the coating from at least one end face of the platform in the region of the overhang zone while leaving the at least first layer of the coating on the same end face in the region of the contact zone.

In particular, the wing can be closed off at both ends by a respective platform, and in particular the said method steps can be carried out on both platforms. The invention is based on the following considerations:
For aerodynamic and mechanical reasons, two adjacent platforms of each two adjacent turbine blades are to be arranged as closely spaced as possible from each other.

If a coating is applied to the wing of a turbine blade on the wing side, for example in order to improve the heat resistance, undesired wetting or coating of the end faces of the platform may also occur depending on the specific coating process. In this case, however, there is usually no controlled application to the face, but the application is irregular and therefore does not have a well-defined density. At certain points of two opposite end surfaces may thereby irregularities of the excess coating touch, resulting in between The two adjacent platforms may form a gap, which is undesirable.

This could now be prevented by preventing any application of the coating on the end faces. Depending on the specific technical design of the application process, however, this is very expensive. For procedural reasons, therefore, at least a partial application The coating can not be prevented on the faces under relatively little effort. However, the complete removal of individual layers of the coating from the end faces of the platform is also associated with a high cost. Moreover, damage to individual wing-side layers of the coating of the platform can thereby occur, in particular at the edges.

An essential, completely surprising finding for the invention is that the platform exhibits a greater thermal expansion in the region of the overhang zone during operation of the turbine than in the region of the contact zone.

This is for a number of reasons: on the one hand, in the overhang zone, the platform is exposed to the high temperatures encountered during operation throughout its areal extent, whereas the platform in the contact zone is not exposed to these temperatures at the points where the wing adjoins. Assuming at a given temperature, a uniform, thermal expansion of all microscopic surface elements of the platform, this means, in simple terms, that in the supernatant zone macroscopically a larger heat-related expansion takes place, since all microscopic surface elements are exposed to the heat, while this for some microscopic surface elements in the contact zone is not the case.

This effect is enhanced by the fact that in the contact zone in the immediate vicinity of the areas where the platform terminates the wing, the thermal expansion of a microscopic surface element of the platform at a predetermined temperature is lower than in other areas. At these points where the platform is connected to the wing or merges into it, the turbine blade, similar to a T-beam, has a particularly high strength. This also affects the thermal expansion in this area.

The said, different thermal expansion of the platform in the region of the contact zone or the supernatant zone is now exploited for the invention in that the at least first layer of the coating is removed from an end face of the platform only in the region of the supernatant zone, while the corresponding layer of the coating is maintained in the region of the contact zone on the end face of the platform. In this case, it is preferable to carry out this procedure on all end faces which, in the blading of a turbine stage, respectively directly oppose one end face of another platform.

During operation of the turbine, a platform now expands more in the area of the overhang zone than in the area of the contact zone, thereby reducing the distance to an adjacent platform. The fact that in the region of the contact zone at the end face of the platform, a layer of the coating is maintained, can now be prevented that due to the stronger thermal expansion in the region of the overhang zone to an adjacent platform in the region of the contact zone forms an excessive gap. The distance between two adjacent platforms can therefore be chosen so that their faces in the region of the overhang zone in operation almost or completely touch, and thus a fluidic escape of hot gas is difficult. In the region of the contact zone of the two adjacent platforms, the escape is made more difficult by the layers of the coating remaining on the end surfaces.

In view of the fact that, in the case of a large number of possible geometries for a turbine blade, the platform merges into the wing in the contact zone to form a concave surface, leaving the at least first layer of the coating on the end face in the region of the contact zone is particularly advantageous. The removal of portions of the coating can result in tangential forces within a layer of the coating. While such tangential Forces in a position of the coating in the supernatant zone can propagate largely unhindered, they can locally have a normal component on a concave surface in the region of the contact zone. This normal force component favors local detachment of the particular layer of coating from the platform in the region of the concave surface and the contact zone. By eliminating the removal of the at least first layer of the coating in the end face in the region of the contact zone, the risk of local detachment of individual layers can thus be considerably reduced.

Preferably, at least the first layer of the coating is applied to the platform on the wing side by spraying. In this case, the specified method is particularly advantageous because it can easily come to a - possibly uncontrolled - wetting the end face of the platform with a layer of the coating at the wing-side spraying the platform with the coating. Thus, the conditions for the procedure are given.

Alternatively, at least the first layer of the coating may be applied to the platform by a dip bath. In particular, in the case of a turbine blade, the wing of each of which is closed off by a platform at one end, it is difficult in this case to prevent a layer of the coating from being applied to the end faces of the platforms. Thus, the conditions for the procedure are met here, too.

Conveniently, a binding layer is applied to the platform on the wing side as the first layer of the coating. Such a binding layer serves to improve the bonding of a further, later to be applied, layer of the coating to the material of the turbine blade. Thus, the turbine blade, in particular on the platform, can be coated with a material which can be optimized for heat and corrosion resistance. The adhesion of this material to the material of the turbine blade must at This optimization should not be considered separately, since the adhesion is ensured by the bonding layer.

If such a bonding layer now forms a first layer of the coating on the platform on the wing side, the specified process is particularly advantageous since it leaves the bonding layer completely intact in the sensitive regions of the contact zone. Thus, there is no danger here that the removal of residues of the bonding layer on the end face could damage the same on the wing side of the platform in the area of the contact zone, which otherwise could lead to reduced adhesion of further layers of the coating in this area.

Conveniently, a superalloy is applied to the platform as the bonding layer. Such a superalloy may in particular be a metal-chromium-aluminum-yttrium compound (MCrAlY), it being possible to use nickel and / or cobalt as the base metal. Superalloys have particularly good properties in terms of their adhesion to materials commonly used for turbine blades.

It proves to be further advantageous if the at least first layer of the coating is removed from the at least one end face of the platform in the region of the projection zone by grinding. This type of removal can be controlled very precisely locally compared to, for example, erosive methods, whereby the risk of undesired damage to layers of the coating can be reduced.

In a further embodiment of the invention, a heat barrier layer is applied to the platform on the wing side as a further layer of the coating. In particular, this may be formed as a ceramic thermal barrier layer. The use of said method is particularly advantageous in this case, as this the risk of damage of layers of the coating which have been applied before the thermal barrier layer, in particular in the sensitive region of the contact zone can be significantly reduced. This has a particularly positive effect on the adhesion of the thermal barrier layer.

The invention further mentions a gas turbine, comprising at least one vane and / or blade, which is coated by the method described above. The advantages stated for the method and its developments can be transferred analogously to the gas turbine.

FIG 1

in einer Schrägansicht eine Plattform einer Turbinenschaufel mit angedeutetem Flügelstumpf,

FIG 2

in einer Draufsicht zwei benachbarte Plattformen von Turbinenschaufeln,

FIG 3

in einem Blockdiagramm den Ablauf eines Verfahrens zur Beschichtung einer Turbinenschaufel, und

FIG 4

in einer Querschnittdarstellung eine Gasturbine.

An embodiment of the invention will be explained in more detail with reference to a drawing. Here are shown schematically in each case:

FIG. 1

in an oblique view, a platform of a turbine blade with indicated stump wing,

FIG. 2

in a plan view, two adjacent platforms of turbine blades,

FIG. 3

in a block diagram the flow of a method for coating a turbine blade, and

FIG. 4

in a cross-sectional view of a gas turbine.

Corresponding parts and sizes are provided in all figures with the same reference numerals.

In FIG. 1 is schematically shown in an oblique view of an end of a turbine blade 1. The turbine blade 1 has a platform 2 and a wing 4, wherein the wing 4 is indicated in this illustration as a wing stump. The region of the platform 2, in which it has contact with the wing 4 or merges into it, is defined here as the contact zone 6. This is characterized by a dotted border. To the contact zone 6 adjacent to the platform 2 in both directions in each case a wing 4 projecting supernatant zone 8a, 8b. The supernatant zones 8a, 8b are each characterized by a dashed border. If a first layer of a coating is applied to the turbine blade 1 by spraying, it is usually not possible to prevent the platform 2 from being sprayed on the side of the platform 2 with parts of the coating. These remnants of the first layer of the coating are to be removed from the end face 10a, 10b in the region of the projection zones 8a, 8b. In contrast, in the region of the contact zone 6, the portions of the coating which have reached the end face 10c when the first layer is applied are left there.

In FIG. 2 are schematically shown in a plan view two adjacent platforms 2 of turbine blade 1. Each of the two platforms 2 in this case has an end face 10 which lies opposite the end face 10 of the respective other platform. If a layer of a coating is now applied to the turbine blade 1 and parts of this layer of the coating also fall on the respective end face 10 of the platform 2, this can lead to the gap 12, by which the two platforms 2 are spaced apart from one another , no longer has the defined width. To counteract this, remnants of the applied layer of the coating are removed in the region of the projection zones 8a, 8b of the two platforms 2 at the end faces 10a, 10b. In the region of the contact zones 6 of each platform 2, in each of which the wing 4 connects to the platform 2, the remains of the layer of the coating on the end face 10c are left in each case.

During operation of a gas turbine with the turbine blades 1, the platforms 2 expand in the transition zones 8a, 8b due to heat stronger than in the contact zones 6. This results in that the gap 12 between two adjacent platforms 2 in the region of the supernatant zones 8a, 8b during the operation has a smaller distance. By leaving residues of a layer of the coating on the end faces 10c in the region of the contact zone 6 can thus be prevented that the gap 12 in this area has too large a width, which could fluidly lead to an undesirable escape of hot gas.

In FIG. 3 1 is a schematic block diagram of the sequence of a method 20 for coating a turbine blade 1. On a turbine blade 1, a first layer 22 of a coating 24 is first applied by spraying 26 on the wing side. The first layer 22 of the coating 24 is a superalloy 28, for example MCrAlY. By spraying 26 of the platform 2 of the turbine blade 1 with the superalloy 28, this is also partly applied to the end face 10 of the platform 2 with.

In a next method step, in the region of the projection zone 8a, 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 zone 8a, the platform is only wing-side after the grinding 30, but no longer coated with the superalloy 28 on the end face 10a. In the region of the contact zone, not shown, the superalloy 28 is retained there on the end face further.

In a further method step, a further layer of the coating 24 is applied on the wing side. This further layer is formed by a ceramic TBC 32. For this TBC 32, the superalloy 28 is fictitious as a bonding layer 34, that is, the superalloy 28 significantly improves the adhesion of the TBC 32 to the blade 1. Against this background, it is particularly advantageous to leave the end face 10 in the region of the contact zone during grinding 30 in order not to damage the first layer 22 of the coating 24 formed by the superalloy 28 in this sensitive area.

In FIG. 4 is shown schematically in a cross-sectional view of a gas turbine 40 with turbine blades 1, which were coated according to the method described above. The turbine blades 1 can in this case be designed both as guide vanes 42 and as rotor blades 44.

Although the invention has been illustrated and described in detail by the preferred embodiment, the invention is not limited by this embodiment. Other variations can be deduced therefrom by those skilled in the art without departing from the scope of the invention.

Claims (8)

1. Method (20) for coating a turbine blade (1), which comprises an aerofoil (4) and at least one platform (2) arranged at one end of the aerofoil (4), wherein the or each platform (2) has a contact zone (6) and at least one two-dimensional projecting zone (8a, 8b), adjoining the contact zone (6), wherein the projecting zone (8a, 8b) can be subjected in its entire two-dimensional extent to the high temperatures occurring during operation, and on the other hand the contact zone (6) cannot be subjected to these temperatures at the locations at which the aerofoil is connected to the platform, and consequently merges with it, wherein the platform has at least one end face (10a, 10b), which during use as intended in the blading of a turbine stage respectively lies directly opposite an end face of another platform, comprising the method steps of:

   - applying on the aerofoil side at least one first layer (22) of a coating (24) to the platform (2), and

   - removing the at least one first layer (22) of the coating (24) from at least one end face (10a, 10b) of the platform (2) in the region of the projecting zone (8a, 8b), while leaving the at least one first layer (22) of the coating (24) on the end face (10c) in the region of the contact zone (6).
2. Method (20) according to Claim 1,
   wherein at least the first layer (22) of the coating (24) is applied to the platform (2) on the aerofoil side by spraying (26) .
3. Method (20) according to Claim 1,
   wherein at least the first layer (22) of the coating (24) is applied to the platform (2) by an immersion bath.
4. Method (20) according to one of the preceding claims,
   wherein a binding layer (34) is applied to the platform (2) as the first layer (22) of the coating (24) on the aerofoil side.
5. Method (20) according to Claim 4,
   wherein a superalloy (28) is applied to the platform (2) as the binding layer (34).
6. Method (20) according to one of the preceding claims,
   wherein at least the first layer (22) of the coating (24) is removed from the at least one end face (10a, 10b) of the platform (2) in the region of the projecting zone (8a, 8b) by grinding.
7. Method (20) according to one of the preceding claims,
   wherein a thermal barrier coating (32) is applied to the platform (2) as a further layer of the coating (24) on the aerofoil side.
8. Gas turbine (40), comprising at least a stationary blade (42) and/or a moving blade (44) which is coated by means of a method according to one of the preceding claims.

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