WO2017148916A1

WO2017148916A1 - Layer system having a coating recess on cooling air holes of turbine blades

Info

Publication number: WO2017148916A1
Application number: PCT/EP2017/054597
Authority: WO
Inventors: Fathi Ahmad; Christian Menke; Christian Voss
Original assignee: Siemens Aktiengesellschaft
Priority date: 2016-03-02
Filing date: 2017-02-28
Publication date: 2017-09-08
Also published as: DE102016203388A1

Abstract

The spallation behavior of a TBC is significantly improved by means of a special recess around cooling air holes (11`, 11``,...), in which an enveloping recess (13`, 13``) and a strip-shaped depression (8), which connects the enveloping recesses to each other, are provided.

Description

Coating system with coating recess on cooling air holes of turbine blades

The invention relates to the recess of a coating around cooling air holes in turbine blades.

The guide and moving blades of the first turbine stages of gas turbines are due to the high thermal load usually with a thermal protection layer (thermal

Barrier Coating = TBC) provided. After coating the turbine blade with the TBC, the previously introduced cooling air holes have reopened and first drilled ^¬ to. Due to the drilling method used, it has so far been unavoidable that microcracks are initiated in the edge regions of the respective openings. These micro-cracks may be the starting point for TBC flaking and should therefore no longer be present after completion of the blade production.

It is therefore an object of the invention to provide a system which prevents the chipping of the TBC around the cooling air holes. The object is achieved by a layer system according to claim 1. An ^¬ .

The subclaims list further advantageous measures which can be combined with one another as desired in order to achieve further advantages.

1, 2 embodiments of the

Invention. As an invention here is an optimal pattern for a TBC

Recess be shown. The goal is still to avoid microcracks in the TBC to eliminate cooling air holes. The pattern can be applied both before applying the TBC by means of a marker, so that the blade is not coated at these points even with a TBC or it is subsequently pasted with a negative mask and then mechanically processed, so that the TBC then only at these points is removed again, for example by

Sandblasting.

The optimal recess consists of the following geometric shapes:

1. Circular surfaces or oval surfaces around the cooling air holes;

2. strip-shaped transverse groove axially behind these surfaces, which connects all circular surfaces or oval surfaces of a row of cooling air.

The figure and the description represent only embodiments of the invention.

1 shows a plan view of an airfoil 4 of a turbine component and a turbine blade 1 - here as an example component.

The airfoil 4 has at the leading edge and on the suction or pressure side cooling air holes 11 11 ^{λ λ} , ..., which in particular in several rows of cooling air 10 10 ^{λ λ} , ...

are arranged.

In addition, the turbine blade 1 has an outer layer 9 on a substrate which is applied to the preferably used nickel- or cobalt-based substrate and has a metallic adhesion promoter layer and / or an outer ceramic layer 9.

To a part 7 (indicated by dashed lines) of the cooling air holes 11 11 ^{λ λ} , so not all cooling holes 11 11 ^{λ λ} , ..., a recess 12 of the outer layer 9 is present. FIG. 2 shows an enlarged view of these cooling air holes 11 11 ^{λ λ} ,... According to FIG. 1.

Around these cooling air holes 11 11 ^{λ λ} , ... around is in each case a circular or oval-shaped recess 13 13 ^{λ λ} , ...

present, which extends around the respective entire cooling air hole 11 11 ^{λ λ} , ....

Furthermore, a strip-shaped recess 8 is present, which is at the recesses 13 13 ^{λ λ of} the cooling air holes

11 11 ^{λ λ} , ... directly connects and the continuous Ausspa ^¬ tion 12 forms.

The strip-shaped recess 8 is preferably rectangular.

In this strip-shaped recess 8 are no

Cooling holes available.

The strip-shaped recess 8 preferably extends only along the cooling air holes 11 11 ^{λ λ} , the

Recesses 13 13 ^{λ λ} , ... have, but not beyond and is fluidically behind the cooling air holes 11 11 ^{λ λ} , ... arranged. The pattern is only in a certain range of

Cooling air holes introduced (half a row of cooling air). The recesses can also be applied to larger parts of the blade. The recess 13 13 ^{λ λ} , 8 can be created by subsequent Bear ^¬ processing or generated again during coating by appropriate abolition.

The advantages are:

• The introduced relief pattern has the potential to reduce or even eliminate TBC flaking while the turbine blade is in operation. Depending on the degree of impact on the spalling behavior, this may reduce unplanned downtime and unplanned blade noise. The pattern shown here has a higher thermal protection function,

· The removal around the cooling air holes removes the microcracks inevitably introduced during drilling. In addition, the transverse groove connecting all circular notches reduces the tendency for the TBC to flake off because it represents a crack propagation barrier.

· Cost savings in the service of gas turbines by

Increased reliability of the gas turbine and thus compliance with guarantees (availability) as well as increased customer benefit and thus strong customer loyalty / potential for new orders.

Claims

claims

1st layer system (1),

at least having:

a substrate with at least one outer layer (9), in particular with an outer ceramic layer (9), wherein the layer system (1) has a plurality of cooling air holes (11 11 ^{λ λ} , ...),

in particular in at least one cooling air row (10

10 ^{λ λ} , ...) arranged and

wherein a plurality of cooling air holes (11 11 ^{λ λ} , ...) at least one common, continuous recess (12) in the

Layer (9) is present,

wherein recess (12) means that there is at least no outer layer (9),

wherein the recess (12) is formed by

a strip-shaped recess (8) and

Circular or oval-shaped recesses (13 13 ^{λ λ} , ...) to the cooling air holes (11, 11 ^{λ λ} , ...),

which are connected to the strip-shaped recess (8), the (8) also has no layer (9).

2. Layer system according to claim 1,

wherein the outer layer (9) is an outer ceramic

Layer (9) represents.

3. Layer system according to one or both of claims 1 or 2,

wherein the outer ceramic layer (9) has a teilstabili ^¬ catalyzed zirconium.

4. Layer system according to one or more of claims 1, 2 or 3,

in which under the outer ceramic layer (9) has a metallic adhesion promoter layer,

especially on the basis NiCoCrAlY is present.

5. Layer system according to claim 4,

in which in the recesses (13 13 ^{λ λ} , 8, 12) no outer ceramic layer (9),

but a metallic primer coating is present.

6. Layer system according to one or more of claims 1, 2, 3, 4 or 5,

in which the recess (12) relates only to a part of cooling air holes (11 11 ^{λ λ} , ...) of a cooling air row (10 10 ^{λ λ} , ...) of a turbine blade.

7. Layer system according to one or more of claims 1, 2, 3, 4, 5 or 6,

in which the strip-shaped recess (8) extends only along the cooling air holes (11 11 ^{λ λ} , ...),

the recesses (13 13 ^{λ λ} , ...) have.

8. Layer system according to one or more of claims 1, 2, 3, 4, 5, 6 or 7,

in which the strip-shaped recess (8) fluidly behind the cooling air holes (11 11 ^{λ λ} , ...) is arranged.

9. Layer system according to one or more of claims 1, 2, 3, 4, 5, 6, 7 or 8,

in which the strip-shaped recess (8) is formed rectangular ^¬ .