Method of Repairing a Component, and a Component

Abstract

A method for the crack repair of a component having a crack on a surface, including excavating a hollow around the crack, wherein the hollow has two recesses and the recesses are arranged at the height of the surface and are open at the height of the surface, and wherein the recesses are arranged around the circumference of the hollow and further including filling the hollow with a braze. A component having a crack on a surface is also provided with a hollow which is excavated around the crack, the hollow has two recesses and the recesses are arranged at the height of the surface and are open at the height of the surface, and the recesses are arranged around the circumference of the hollow, wherein the hollow is filled with the braze.

Description

FIELD OF INVENTION

The invention relates to a method for repairing a component and to a component.

BACKGROUND OF INVENTION

Newly produced components, for example cast components, or components after use, often have cracks which are repaired so that the component can be used again.

According to the prior art, material is excavated around the crack and it is filled with a braze or closed by welding.

So-called coupon brazing is also known from EP 0 868 253 B1.

The methods mentioned above, however, often cannot sufficiently prevent crack growth during reuse of the component.

SUMMARY OF INVENTION

It is therefore an object of the invention to overcome the problem mentioned above.

The object is achieved by a method and a component as claimed in the independent claims.

BRIEF DESCRIPTION OF THE DRAWINGS

The invention will be explained in more detail with the aid of the figures.

The dependent claims list other advantageous measures, which may advantageously be combined with one another in any desired way.

FIG. 1 shows a component surface with a crack,

FIG. 2 shows a sectional representation of FIG. 1 along the line II-II,

FIG. 3 shows a sectional representation of a hollow,

FIG. 4 shows a surface of a component with a hollow,

FIG. 5 shows the negative of the excavated crack surface (hollow),

FIG. 6 shows the use of rods in the further repair of the component,

FIG. 7 shows sectional representation of FIG. 6 along the line VII-VII,

FIGS. 8, 9, 10, 11 show other exemplary embodiments of the invention,

FIGS. 12, 13 show another possibility for arranging rods in the hollow,

FIG. 14 shows a gas turbine,

FIG. 15 shows a turbine blade in perspective and

FIG. 16 shows a combustion chamber in perspective.

DETAILED DESCRIPTION OF INVENTION

FIG. 1 shows a surface 4 of a component 1, which has a crack 7 with a length 1 along a crack direction 8. The crack 7 may also extend in a curve, although in this case it also has an (averaged) direction 8.

The component 1 is preferably a turbine blade 120, 130 (FIG. 14) or a combustion chamber element 155 (FIG. 15) of a turbine, for example an aircraft turbine, preferably a gas turbine 100 (FIG. 13).

Particularly in the case of turbine components, a substrate 10 of the component 1 consists of a nickel- or cobalt-based superalloy.

The method of crack repair is not however restricted to such components, rather it also encompasses all components having cracks, cavities, indentations which are repaired by means of brazing or welding, and also those which consist of other materials.

FIG. 2 shows a sectional representation of FIG. 1, from which it may be seen that the crack 7 has a crack depth a in the substrate 10 of the component 1, which has a thickness or wall thickness t (a<<t).

The crack 7 according to FIGS. 1, 2 is excavated as in the prior art, so as to create a hollow 11 which according to the prior art comprises an indentation which is rectangular (FIGS. 3, 4, 5) or triangular (FIGS. 8, 9) in cross section. Other cross-sectional shapes may be envisaged. The hollow 11 is preferably of cuboid shape.

However, the hollow 11 additionally comprises at least three recesses 13, 13′, 16, 16′, 19, 19′.

FIG. 4 shows a plan view of such a hollow 11.

The recesses 13, 13′, . . . may be arranged arbitrarily along the direction 8 around the circumference of the hollow 11. The recesses 13, 13′, . . . are preferably arranged lying directly opposite in pairs (FIGS. 4, 5, 6, 9, 10).

The distances between the recesses 13, 13′, 16, 16″, . . . , and in particular also from the start or end 17 of the hollow 11, are preferably equidistant (FIGS. 5, 6). In the exemplary embodiment according to FIG. 4, the recesses 13, 13′ are present at the end 17 of the hollow 11.

FIG. 5 represents the negative of the excavated inner surface of the hollow 11 with recesses 13, 13′, . . .

Here the recesses 13, 13′ are not arranged at the end 17 of the hollow 11, rather they lie at a distance from the end 17 of the hollow 11 along the direction 8.

The excavated hollow 11 has a height a′, which is equal or essentially corresponds to the crack depth a according to FIG. 3 (a≦a′). The hollow 11 has a length l′ along the crack direction 8, which is equal or essentially corresponds to the length l of the crack 7 (FIG. 1) (l≦l′).

Furthermore, the hollow 11 (without recesses) has a width b which is wide enough so that the crack 7 and attacked surfaces on the crack surface have been removed.

The recesses 13, 16, 19, which preferably have recesses 13′, 16′, 19′ lying opposite, in particular perpendicularly to the crack direction 8, substantially represent cuboids transversely to the direction 8 (see dashed indication) which have a length b+greater than the width b of the hollow 11.

The recesses 13, 13′, . . . are for example of cuboid (or cubic) shape here and have a length f along the direction 8, a depth g in the direction of the depth a′ and a width e in the direction of the width b. The width e of the recess is preferably much less than the width b of the hollow 11, the length f is much less than the length l′ of the hollow 11 and the depth g is preferably much less than the depth a′ of the hollow 11.

In FIG. 5, the recesses 13, 13′, . . . are of cubic or cuboid shape. The recesses 13, 13′, . . . may likewise have other shapes (round or triangular in cross section).

After the crack has been excavated according to FIG. 5, rods 22′, 22″, 22′″ are placed into the recesses 13, 13′, . . . (FIG. 6), the melting temperature of these rods being higher than the melting temperature of the filler material, which constitutes in particular a braze 25, with which the hollow 11 to be filled is then also filled (FIG. 7). The hollow 11 may also be closed by welding.

The recesses 13, 13′ are preferably configured so that the rods 22′, 22″, 22′″ can be put into them from above, i.e. the recesses 13, 13′, . . . are open at the height of the surface 4, and they do not protrude from the hollow 11 beyond the surface 4. The depth g of the recesses 13, 13′, . . . is preferably configured here (FIG. 5) so that the rod 22′, 22″, 22′″ is arranged close to the surface 4 of the component 1.

The rods 22′, 22″ are preferably arranged in a plane, i.e. at one height. The height of the rods 22′, 22″ between one another, i.e. the depth g of the respective recesses 13, 13′, . . . , may also be selected arbitrarily so that there is a larger distance from the surface 4.

The material of the rods 22′, 22″, 22′″ preferably consists of a ceramic material.

The effect of the rods is that a crack, which is often formed on the surface 4 and propagates through the braze, meets the ceramic rods 22′, 22″, 22′″ where it is prevented from growing further in them owing to the greater toughness of the rods 22′, 22″, 22′″.

The recesses 13, 13′, 16, 16′, 19, 19′ may likewise extend over the entire height a′ of the hollow 11, as is represented in FIG. 10. Plates instead of rods are then preferably inserted into these recesses 13, 13′, . . . , the plates extending over the entire width b+. The plates may have the height a′, although they may also be of smaller size.

Instead of plates, fine- or coarse-latticed meshes may also be inserted so that the filler material or the braze can enclose the mesh during the brazing and the mesh acts as crack prevention over the entire depth of the hollow 11.

Fiber bundles may also be inserted into the hollow 11 instead of the rods 22′, 22″, 22′″, and fiber mats may also be inserted instead of the plates.

A coupon may also be brazed or welded into the hollow 11, the coupon having a shape according to FIG. 5 while having smaller dimensions than the hollow 11, so that it can be fitted into it.

FIG. 11 shows another exemplary embodiment of an arrangement of the rods 22′, 22″, 22′″ in the hollow 11.

The rods 22′, 22″, . . . do not extend perpendicularly to the crack direction 8 or to the longitudinal direction of the hollow 11 here, rather they are arranged obliquely. The rods 22″, 22′″ may likewise cross over, i.e. either they are arranged at different heights in the hollow 11 or a cross in the form of an “X” is placed into the recesses 16, 16′, 19, 19′, in which case the rods 22″, 22′″ may form one part.

Further arrangement possibilities for the rods 22′, 22″ with respect to one another, or the combination of rods and fiber mats (FIG. 10), may be envisaged.

Instead of the recesses, which adjoin the surface 4 of the component 1, there may also be recesses 13, 13′, . . . below the surface 4 on the inner surfaces of the hollow 11. The rods 22, . . . must then be flexible enough to be bent (FIG. 12) so that they can be inserted into the hollow 11, in which case the rods must have a length >b so that their two ends rest in the indentations 13, 13′ (FIG. 13).

FIG. 14 shows a gas turbine 100 by way of example in a partial longitudinal section.

The gas turbine 100 internally comprises a rotor 103, which will also be referred to as the turbine rotor, mounted so as to rotate about a rotation axis 102 and having a shaft 101.

Successively along the rotor 103, there are an intake manifold 104, a compressor 105, an e.g. toroidal combustion chamber 110, in particular a ring combustion chamber, having a plurality of burners 107 arranged coaxially, a turbine 108 and the exhaust manifold 109. The ring combustion chamber 110 communicates with an e.g. annular hot gas channel 111. There, for example, four successively connected turbine stages 112 form the turbine 108. Each turbine stage 112 is formed for example by two blade rings. As seen in the flow direction of a working medium 113, a guide vane row 115 is followed in the hot gas channel 111 by a row 125 formed by rotor blades 120.

The guide vanes 130 are fastened on an inner housing 138 of a stator 143 while the rotor blades 120 of a row 125 are fitted on the rotor 103, for example by means of a turbine disk 133. Coupled to the rotor 103, there is a generator or a work engine (not shown).

During operation of the gas turbine 100, air 135 is taken in and compressed by the compressor 105 through the intake manifold 104. The compressed air provided at the end of the compressor 105 on the turbine side is delivered to the burners 107 and mixed there with a fuel. The mixture is then burnt to form the working medium 113 in the combustion chamber 110. From there, the working medium 113 flows along the hot gas channel 111 past the guide vanes 130 and the rotor blades 120. At the rotor blades 120, the working medium 113 expands by imparting momentum, so that the rotor blades 120 drive the rotor 103 and the work engine coupled to it.

During operation of the gas turbine 100, the components exposed to the hot working medium 113 experience thermal loads. Apart from the heat shield elements lining the ring combustion chamber 110, the guide vanes 130 and rotor blades 120 of the first turbine stage 112, as seen in the flow direction of the working medium 113, are heated the most.

In order to withstand the temperatures prevailing there, they may be cooled by means of a coolant. Substrates of the components may likewise comprise a directional structure, i.e. they are monocrystalline (SX structure) or comprise only longitudinally directed grains (DS structure). Iron-, nickel- or cobalt-based superalloys used as material for the components, in particular for the turbine blades 120, 130 and components of the combustion chamber 110. Such superalloys are known for example from EP 1 204 776 B1, EP 1 306 454, EP 1 319 729 A1, WO 99/67435 or WO 00/44949; with respect to the chemical composition of the alloys, these documents are part of the disclosure.

The guide vane 130 comprises a guide vane root (not shown here) facing the inner housing 138 of the turbine 108, and a guide vane head lying opposite the guide vane root. The guide vane head faces the rotor 103 and is fixed on a fastening ring 140 of the stator 143.

FIG. 15 shows a perspective view of a rotor blade 120 or guide vane 130 of a turbomachine, which extends along a longitudinal axis 121.

The turbomachine may be a gas turbine of an aircraft or of a power plant for electricity generation, a steam turbine or a compressor.

The blade 120, 130 comprises, successively along the longitudinal axis 121, a fastening region 400, a blade platform 403 adjacent thereto as well as a blade surface 406 and a blade tip 415. As a guide vane 130, the vane 130 may have a further platform (not shown) at its vane tip 415.

A blade root 183 which is used to fasten the rotor blades 120, 130 on a shaft or a disk (not shown) is formed in the fastening region 400. The blade root 183 is configured, for example, as a hammerhead. Other configurations as a fir-tree or dovetail root are possible. The blade 120, 130 comprises a leading edge 409 and a trailing edge 412 for a medium which flows past the blade surface 406.

In conventional blades 120, 130, for example solid metallic materials, in particular superalloys, are used in all regions 400, 403, 406 of the blade 120, 130. Such superalloys are known for example from EP 1 204 776 B1, EP 1 306 454, EP 1 319 729 A1, WO 99/67435 or WO 00/44949; with respect to the chemical composition of the alloy, these documents are part of the disclosure. The blades 120, 130 may in this case be manufactured by a casting method, also by means of directional solidification, by a forging method, by a machining method or combinations thereof.

Workpieces with a monocrystalline structure or structures are used as components for machines which during operation are exposed to heavy mechanical, thermal and/or chemical loads. Such monocrystalline workpieces are manufactured, for example, by directional solidification from the melts. These are casting methods in which the liquid metal alloy is solidified to form a monocrystalline structure, i.e. to form the monocrystalline workpiece, or is directionally solidified. Dendritic crystals are in this case aligned along the heat flux and form either a rod crystalline grain structure (columnar, i.e. grains which extend over the entire length of the workpiece and in this case, according to general terminology usage, are referred to as directionally solidified) or a monocrystalline structure, i.e. the entire workpiece consists of a single crystal. It is necessary to avoid the transition to globulitic (polycrystalline) solidification in these methods, since nondirectional growth will necessarily form transverse and longitudinal grain boundaries which negate the beneficial properties of the directionally solidified or monocrystalline component. When directionally solidified structures are referred to in general, this is intended to mean both single crystals which have no grain boundaries or at most small-angle grain boundaries, and also rod crystal structures which, although they do have grain boundaries extending in the longitudinal direction, do not have any transverse grain boundaries. These latter crystalline structures are also referred to as directionally solidified structures. Such methods are known from U.S. Pat. No. 6,024,792 and EP 0 892 090 A1; with respect to the solidification method, these documents are part of the disclosure.

The blades 120, 130 may likewise have coatings against corrosion or oxidation, for example (MCrAlX; M is at least one element from the group iron (Fe), cobalt (Co), nickel (Ni), X is an active element and stands for yttrium (Y) and/or silicon and/or at least one rare earth element, or hafnium (Hf)). Such alloys are known from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1 which, with respect to the chemical composition of the alloy, are intended to be part of this disclosure. The density may preferably be 95% of the theoretical density. A protective aluminum oxide layer (TGO=thermal grown oxide layer) is formed on the MCrAlX layer (as an interlayer or as the outermost layer).

On the MCrAlX, there may furthermore be a thermal barrier layer, which is preferably the outermost layer and consists for example of ZrO₂, Y₂O₃—ZrO₂, i.e. it is not stabilized or is partially or fully stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide. The thermal barrier layer covers the entire MCrAlX layer. Rod-shaped grains are produced in the thermal barrier layer by suitable coating methods, for example electron beam deposition (EB-PVD). Other coating methods may be envisaged, for example atmospheric plasma spraying (APS), LPPS, VPS or CVD. The thermal barrier layer may comprise porous, micro- or macro-cracked grains for better thermal shock resistance. The thermal barrier layer is thus preferably more porous than the MCrAlX layer.

The blade 120, 130 may be designed to be hollow or solid. If the blade 120, 130 is intended to be cooled, it will be hollow and optionally also comprise film cooling holes 418 (indicated by dashes).

FIG. 16 shows a combustion chamber 110 of a gas turbine 100. The combustion chamber 110 is designed for example as a so-called ring combustion chamber in which a multiplicity of burners 107, which produce flames 156 and are arranged in the circumferential direction around a rotation axis 102, open into a common combustion chamber space 154. To this end, the combustion chamber 110 as a whole is designed as an annular structure which is positioned around the rotation axis 102.

In order to achieve a comparatively high efficiency, the combustion chamber 110 is designed for a relatively high temperature of the working medium M, i.e. about 1000° C. to 1600° C. In order to permit a comparatively long operating time even under these operating parameters which are unfavorable for the materials, the combustion chamber wall 153 is provided with an inner lining formed by heat shield elements 155 on its side facing the working medium M.

Owing to the high temperatures inside the combustion chamber 110, a cooling system may also be provided for the heat shield elements 155 or for their retaining elements. The heat shield elements 155 are then hollow, for example, and optionally also have film cooling holes (not shown) opening into the combustion chamber space 154.

Each heat shield element 155 made of an alloy is equipped with a particularly heat-resistant protective layer (MCrAlX layer and/or ceramic coating) on the working medium side, or is made of refractory material (solid ceramic blocks). These protective layers may be similar to the turbine blades, i.e. for example MCrAlX means: M is at least one element from the group iron (Fe), cobalt (Co), nickel (Ni), X is an active element and stands for yttrium (Y) and/or silicon and/or at least one rare earth element, or hafnium (Hf). Such alloys are known from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1 which, with respect to the chemical composition of the alloy, are intended to be part of this disclosure.

On the MCrAlX, there may furthermore be an e.g. ceramic thermal barrier layer which consists for example of ZrO₂, Y₂O₃—ZrO₂, i.e. it is not stabilized or is partially or fully stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide.

Rod-shaped grains are produced in the thermal barrier layer by suitable coating methods, for example electron beam deposition (EB-PVD). Other coating methods may be envisaged, for example atmospheric plasma spraying (APS), LPPS, VPS or CVD. The thermal barrier layer may comprise porous, micro- or macro-cracked grains for better thermal shock resistance.

Refurbishment means that turbine blades 120, 130 and heat shield elements 155 may need to have protective layers taken off (for example by sandblasting) after their use. The corrosion and/or oxidation layers or products are then removed. Optionally, cracks in the turbine blade 120, 130 or the heat shield element 155 are also repaired by the method. The turbine blades 120, 130 or heat shield elements 155 are then recoated and the turbine blades 120, 130 or the heat shield elements 155 are used again.

Claims

1.-13. (canceled)

14. A method for the crack repair of a component having a crack on a surface, comprising: excavating a hollow around the crack, the hollow has two recesses and the recesses are arranged at a height of the surface and are open at the height of the surface, and the recesses are arranged around the circumference of the hollow; andfilling the hollow with a braze.

15. The method as claimed in claim 14, wherein the recesses are arranged in pairs on opposite sides of the hollow.

16. The method as claimed in claim 14, wherein the hollow is cuboid.

17. The method as claimed in claim 15, wherein the hollow is cuboid.

18. The method as claimed in claim 14, wherein the hollow has a longitudinal direction parallel to the crack direction of the crack

19. The method as claimed in claim 14, wherein the recesses extend transversely to a crack direction of the crack or a longitudinal direction of the hollow and accordingly the hollow has a width at the height of the surface greater than a width of the hollow without a recess.

20. The method as claimed in claim 19, wherein the recesses extend perpendicularly to a crack direction of the crack or a longitudinal direction of the hollow.

21. The method as claimed in claim 14, wherein the recesses form a cuboid.

22. The method as claimed in claim 14, further comprising: arranging rods on the component, wherein the recesses are in such a shape that the rods are arranged in the vicinity of the surface of the component, but do not protrude beyond the surface of the component.

23. The method as claimed in claim 21, further comprising: arranging rods on the component, wherein the recesses are in such a shape that the rods are arranged in the vicinity of the surface of the component, but do not protrude beyond the surface of the component.

24. The method as claimed in claim 14, further comprising: arranging rods in the recesses of the hollow; andfilling the hollow with a filler material, wherein the melting temperature of the rods is higher than the melting temperature of the filler material.

25. The method as claimed in claim 24, wherein the hollow is filled with a braze.

26. The method as claimed in claim 22, wherein the rods are made of ceramic.

27. The method as claimed in claim 24, wherein the rods are made of ceramic.

28. The method as claimed in claim 14, further comprising: fastening a coupon into the hollow, wherein the coupon has a shape of the hollow with recesses.

29. The method as claimed in claim 28, wherein the coupon is brazed in the hollow.

30. A component, comprising: a crack on a surface;a hollow which is excavated around the crack, the hollow has two recesses and the recesses are arranged at the height of the surface and are open at the height of the surface, and the recesses are arranged around the circumference of the hollow; anda braze, wherein the hollow is filled with the braze.

31. The component as claimed in claim 30, wherein the recesses are arranged in pairs on opposite sides of the hollow.

32. The component as claimed in claim 30, wherein the hollow is cuboid.

33. The component as claimed in claim 30, wherein rods are arranged on the component, wherein the recesses are in such a shape that the rods are arranged in the vicinity of the surface of the component, but do not protrude beyond the surface of the component.