This application is the US National Stage of International Application No. PCT/EP2010/054322, filed Mar. 31, 2010 and claims the benefit thereof. The International Application claims the benefits of German application No. 10 2009 016 260.7 DE filed Apr. 3, 2009. All of the applications are incorporated by reference herein in their entirety.
The invention relates to a process for filling a recess of a component by welding and to a component.
When repairing components by welding, it is often the case that recesses are also filled. These recesses are produced by the excavation of a damaged region which has arisen during operation of the component. For refurbishment, it is necessary to add material so as to achieve the geometry of the component and also a sufficient strength of the component. Depending on the weld filler, binding defects and cracks may repeatedly occur in the added material.
It is therefore an object of the invention to solve the abovementioned problem.
The problem is solved by a process as claimed in the claims and by a component as claimed in the claims.
The dependent claims list further advantageous measures which can be combined with one another, as desired, in order to achieve further advantages.
The figures and the description represent only exemplary embodiments of the invention.
The component 1 is preferably a turbine blade or vane 120, 130 and preferably consists of a nickel-based or cobalt-based superalloy (
The recess 4 is intended to be filled with welding material. The recess 4 is delimited by a contour 16 (a closed line) with respect to the outer surface 13.
The recess 4 has flanks 28, which preferably run obliquely and not perpendicularly with respect to the surface 13 close to the contour 16 (
The recess 4 is preferably filled by build-up welding.
It is preferable to use a laser process as the welding process.
A welding layer I, II, III (
A main direction 25 of the welding tracks 10′, 10″, 10′″, . . . represents the longest extent 11 of a welding track 10′, 10″, 10′″ and is shown as an arrow in
However, the welding tracks 10′, 10″, 10′″, . . . overshoot the contour 16 of the recess 4 and therefore in part reach the surface 13 (
The cross section of such layers I, II, III laid one above another is shown in
The hump 22 thereby produced (
As a result of the deliberate overshooting, i.e. as a result of additional welding material in the region of the surface 13, good welding results are achieved and no cracks foam in the finish-welded component 1, 120, 130, 155. It is preferable that merely the welding material above the surface 13 has to be removed.
This first outer welding track 6 can lie within the contour 16 (
A further welding track 7 can preferably be laid, likewise corresponding to the outer contour of the recess 4 and lying within the first welding track 6 (
A meandering progression of the welding track 10′, 10″, . . . is then selected within the contour 6, 7 (
Similarly, the longest part 11 of the zigzag curve can run parallel to the longest orientation of the recess 16 (
Recesses do not necessarily have to be filled. It is likewise possible for material to be applied areally to each surface (
The recess 4 is preferably covered completely by a first layer I of welding tracks (
A second welding layer II is then applied, the end of which likewise protrudes beyond the surface 13.
The second welding layer II preferably covers the first welding layer I completely. This provision of layers one above another is continued until a last layer III preferably lies completely over a surface 13.
Here, a first welding layer I is laid by a plurality of welding tracks (10′, 10″, 10′″) with a main direction 25 parallel to the plane of the drawing (the orientation of 25 is arbitrary). The main direction 25 is the longest extent 11 of a welding track 10′, 10″, . . . in the case of a meandering formation (
The second welding layer II is laid with welding tracks in a main direction 25 which is different, preferably perpendicular, to the main direction 25 of the welding layer I, i.e. from the plane of the drawing, a main direction of the welding tracks 10′, 10″, . . . of the third welding layer III preferably running, in turn, like the first welding layer I.
Procedures before the heat treatment of a weld seam 28 are shown on the left-hand side of
Both the left-hand side of
The last step is then the heat treatment (HT) with the weld seam 28, which is conventional depending on the material and the component.
In the interior, the gas turbine 100 has a rotor 103 with a shaft 101 which is mounted such that it can rotate about an axis of rotation 102 and is also referred to as the turbine rotor.
An intake housing 104, a compressor 105, a, for example, toroidal combustion chamber 110, in particular an annular combustion chamber, with a plurality of coaxially arranged burners 107, a turbine 108 and the exhaust-gas housing 109 follow one another along the rotor 103.
The annular combustion chamber 110 is in communication with a, for example, annular hot-gas passage 111, where, by way of example, four successive turbine stages 112 form the turbine 108.
Each turbine stage 112 is formed, for example, from two blade or vane rings. As seen in the direction of flow of a working medium 113, in the hot-gas passage 111 a row of guide vanes 115 is followed by a row 125 formed from rotor blades 120.
The guide vanes 130 are secured to an inner housing 138 of a stator 143, whereas the rotor blades 120 of a row 125 are fitted to the rotor 103 for example by means of a turbine disk 133.
A generator (not shown) is coupled to the rotor 103.
While the gas turbine 100 is operating, the compressor 105 sucks in air 135 through the intake housing 104 and compresses it. The compressed air provided at the turbine-side end of the compressor 105 is passed to the burners 107, where it is mixed with a fuel. The mix is then burnt in the combustion chamber 110, forming the working medium 113. From there, the working medium 113 flows along the hot-gas passage 111 past the guide vanes 130 and the rotor blades 120. The working medium 113 is expanded at the rotor blades 120, transferring its momentum, so that the rotor blades 120 drive the rotor 103 and the latter in turn drives the generator coupled to it.
While the gas turbine 100 is operating, the components which are exposed to the hot working medium 113 are subject to thermal stresses. The guide vanes 130 and rotor blades 120 of the first turbine stage 112, as seen in the direction of flow of the working medium 113, together with the heat shield elements which line the annular combustion chamber 110, are subject to the highest thermal stresses.
To be able to withstand the temperatures which prevail there, they may be cooled by means of a coolant.
Substrates of the components may likewise have a directional structure, i.e. they are in single-crystal faun (SX structure) or have only longitudinally oriented grains (DS structure).
By way of example, iron-based, nickel-based or cobalt-based superalloys are used as material for the components, in particular for the turbine blade or vane 120, 130 and components of the combustion chamber 110.
Superalloys of this type 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.
The blades or vanes 120, 130 may likewise have coatings protecting against corrosion (MCrAlX; M is at least one element selected from the group consisting of iron (Fe), cobalt (Co), nickel (Ni), X is an active element and stands for yttrium (Y) and/or silicon, scandium (Sc) and/or at least one rare earth element, or hafnium). Alloys of this type are known from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1.
It is also possible for a thermal barrier coating to be present on the MCrAlX, consisting for example of ZrO2, Y2O3—ZrO2, i.e. unstabilized, partially stabilized or fully stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide.
Columnar grains are produced in the thermal barrier coating by suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD).
The guide vane 130 has a guide vane root (not shown here), which faces the inner housing 138 of the turbine 108, and a guide vane head which is at the opposite end from the guide vane root. The guide vane head faces the rotor 103 and is fixed to a securing ring 140 of the stator 143.
The turbomachine may be a gas turbine of an aircraft or of a power plant for generating electricity, a steam turbine or a compressor.
The blade or vane 120, 130 has, in succession along the longitudinal axis 121, a securing region 400, an adjoining blade or vane platform 403 and a main blade or vane part 406 and a blade or vane tip 415.
As a guide vane 130, the vane 130 may have a further platform (not shown) at its vane tip 415.
The blade or vane root 183 is designed, for example, in hammerhead form. Other configurations, such as a fir-tree or dovetail root, are possible.
The blade or vane 120, 130 has a leading edge 409 and a trailing edge 412 for a medium which flows past the main blade or vane part 406.
In the case of conventional blades or vanes 120, 130, by way of example solid metallic materials, in particular superalloys, are used in all regions 400, 403, 406 of the blade or vane 120, 130.
Superalloys of this type 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.
The blade or vane 120, 130 may in this case be produced by a casting process, by means of directional solidification, by a forging process, by a milling process or combinations thereof.
Workpieces with a single-crystal structure or structures are used as components for machines which, in operation, are exposed to high mechanical, thermal and/or chemical stresses.
Single-crystal workpieces of this type are produced, for example, by directional solidification from the melt. This involves casting processes in which the liquid metallic alloy solidifies to form the single-crystal structure, i.e. the single-crystal workpiece, or solidifies directionally.
In this case, dendritic crystals are oriented along the direction of heat flow and form either a columnar crystalline grain structure (i.e. grains which run over the entire length of the workpiece and are referred to here, in accordance with the language customarily used, as directionally solidified) or a single-crystal structure, i.e. the entire workpiece consists of one single crystal. In these processes, a transition to globular (polycrystalline) solidification needs to be avoided, since non-directional growth inevitably forms transverse and longitudinal grain boundaries, which negate the favorable properties of the directionally solidified or single-crystal component.
Where the text refers in general terms to directionally solidified microstructures, this is to be understood as meaning both single crystals, which do not have any grain boundaries or at most have small-angle grain boundaries, and columnar crystal structures, which do have grain boundaries running in the longitudinal direction but do not have any transverse grain boundaries. This second form of crystalline structures is also described as directionally solidified microstructures (directionally solidified structures).
Processes of this type are known from U.S. Pat. No. 6,024,792 and EP 0 892 090 A1.
The blades or vanes 120, 130 may likewise have coatings protecting against corrosion or oxidation e.g. (MCrAlX; M is at least one element selected from the group consisting of 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)). Alloys of this type are known from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1.
The density is preferably 95% of the theoretical density.
A protective aluminum oxide layer (TGO=thermally grown oxide layer) is formed on the MCrAlX layer (as an intermediate layer or as the outermost layer).
The layer preferably has a composition Co-30Ni-28Cr-8Al-0.6Y-0.7Si or Co-28Ni-24Cr-10Al-0.6Y. In addition to these cobalt-based protective coatings, it is also preferable to use nickel-based protective layers, such as Ni-10Cr-12Al-0.6Y-3Re or Ni-12Co-21Cr-11Al-0.4Y-2Re or Ni-25Co-17Cr-10Al-0.4Y-1.5Re.
It is also possible for a thermal barrier coating, which is preferably the outermost layer and consists for example of ZrO2, Y2O3—ZrO2, i.e. unstabilized, partially stabilized or fully stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide, to be present on the MCrAlX.
The thermal barrier coating covers the entire MCrAlX layer. Columnar grains are produced in the thermal barrier coating by suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD).
Other coating processes are possible, for example atmospheric plasma spraying (APS), LPPS, VPS or CVD. The thermal barrier coating may include grains that are porous or have micro-cracks or macro-cracks, in order to improve the resistance to thermal shocks. The thermal barrier coating is therefore preferably more porous than the MCrAlX layer.
Refurbishment means that after they have been used, protective layers may have to be removed from components 120, 130 (e.g. by sand-blasting). Then, the corrosion and/or oxidation layers and products are removed. If appropriate, cracks in the component 120, 130 are also repaired. This is followed by recoating of the component 120, 130, after which the component 120, 130 can be reused.
The blade or vane 120, 130 may be hollow or solid in form.
If the blade or vane 120, 130 is to be cooled, it is hollow and may also have film-cooling holes 418 (indicated by dashed lines).
To achieve a relatively high efficiency, the combustion chamber 110 is designed for a relatively high temperature of the working medium M of approximately 1000° C. to 1600° C. To allow a relatively long service life even with these operating parameters, which are unfavorable for the materials, the combustion chamber wall 153 is provided, on its side which faces the working medium M, with an inner lining formed from heat shield elements 155.
On the working medium side, each heat shield element 155 made from an alloy is equipped with a particularly heat-resistant protective layer (MCrAlX layer and/or ceramic coating) or is made from material that is able to withstand high temperatures (solid ceramic bricks).
These protective layers may be similar to the turbine blades or vanes, i.e. for example MCrAlX: M is at least one element selected from the group consisting of 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). Alloys of this type are known from EP 0 486 489 B1, EP 0 786 017 B1, EP 0 412 397 B1 or EP 1 306 454 A1.
It is also possible for a, for example, ceramic thermal barrier coating to be present on the MCrAlX, consisting for example of ZrO2, Y2O3—ZrO2, i.e. unstabilized, partially stabilized or fully stabilized by yttrium oxide and/or calcium oxide and/or magnesium oxide.
Columnar grains are produced in the thermal barrier coating by suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD).
Other coating processes are possible, e.g. atmospheric plasma spraying (APS), LPPS, VPS or CVD. The thermal barrier coating may include grains that are porous or have micro-cracks or macro-cracks, in order to improve the resistance to thermal shocks.
Refurbishment means that after they have been used, protective layers may have to be removed from heat shield elements 155 (e.g. by sand-blasting). Then, the corrosion and/or oxidation layers and products are removed. If appropriate, cracks in the heat shield element 155 are also repaired. This is followed by recoating of the heat shield elements 155, after which the heat shield elements 155 can be reused.
Moreover, a cooling system may be provided for the heat shield elements 155 and/or their holding elements, on account of the high temperatures in the interior of the combustion chamber 110. The heat shield elements 155 are then, for example, hollow and may also have cooling holes (not shown) opening out into the combustion chamber space 154.
Number | Date | Country | Kind |
---|---|---|---|
10 2009 016 260 | Apr 2009 | DE | national |
Filing Document | Filing Date | Country | Kind | 371c Date |
---|---|---|---|---|
PCT/EP2010/054322 | 3/31/2010 | WO | 00 | 1/16/2012 |
Publishing Document | Publishing Date | Country | Kind |
---|---|---|---|
WO2010/112553 | 10/7/2010 | WO | A |
Number | Name | Date | Kind |
---|---|---|---|
5060842 | Qureshi et al. | Oct 1991 | A |
6024792 | Kurz et al. | Feb 2000 | A |
6436553 | Stadler et al. | Aug 2002 | B1 |
20020125216 | Sauron et al. | Sep 2002 | A1 |
20020130112 | Manasas et al. | Sep 2002 | A1 |
20040191064 | Guo | Sep 2004 | A1 |
20050109818 | Katayama | May 2005 | A1 |
20060193612 | Bouet et al. | Aug 2006 | A1 |
20070119830 | Meier | May 2007 | A1 |
Number | Date | Country |
---|---|---|
1322892 | Nov 2001 | CN |
1883872 | Dec 2006 | CN |
69420453 | Apr 2000 | DE |
102007034242 | Apr 2009 | DE |
0486489 | Nov 1994 | EP |
0412397 | Mar 1998 | EP |
0836904 | Apr 1998 | EP |
0892090 | Jan 1999 | EP |
0786017 | Mar 1999 | EP |
1306454 | May 2003 | EP |
1319729 | Jun 2003 | EP |
1204776 | Jun 2004 | EP |
63194884 | Aug 1988 | JP |
63224890 | Sep 1988 | JP |
7241692 | Sep 1995 | JP |
8001332 | Jan 1996 | JP |
H08015481 | Jan 1996 | JP |
8323473 | Dec 1996 | JP |
3164137 | May 2001 | JP |
2001287062 | Oct 2001 | JP |
3272853 | Apr 2002 | JP |
2008264841 | Nov 2008 | JP |
2109611 | Apr 1998 | RU |
2288082 | Nov 2006 | RU |
901304 | Jan 1982 | SU |
1122718 | Nov 1984 | SU |
WO 9967435 | Dec 1999 | WO |
WO 0044949 | Aug 2000 | WO |
Number | Date | Country | |
---|---|---|---|
20120103950 A1 | May 2012 | US |