COMPONENT WITH DIAGONALLY EXTENDING RECESSES IN THE SURFACE AND PROCESS FOR OPERATING A TURBINE

Abstract

The invention is related to a component of a turbine machine with diagonally extending recesses in the surface of the component. Thermally stressed components of a turbine machine can be protected against excessive heat input by active cooling or by applying thermal insulation layers. The invention increases the effect by providing diagonally extending slots in the surface of the component.

Description

CROSS REFERENCE TO RELATED APPLICATIONS

This application is the US National Stage of International Application No. PCT/EP2007/061609 filed Nov. 29, 2007 and claims the benefit thereof. The International Application claims the benefits of European application No. 070000189.6 filed Jan. 5, 2007, both of the applications are incorporated by reference herein in their entirety.

FIELD OF THE INVENTION

The invention relates to a component with diagonally extending recesses on the surface, and to a process for operating a turbine.

BACKGROUND OF THE INVENTION

During use, components having a medium flowing over or around them, such as turbomachines (for example gas turbines), should not exceed certain temperatures and have to be protected against excessive heat input and/or have to be cooled.

In the case of gas turbines, this is done by applying ceramic thermal barrier coatings which, in particular, have a porous design. In addition to the use of porous thermal barrier coatings, film cooling is also known in the case of gas turbine blades or vanes.

U.S. Pat. No. 6,703,137 B2 discloses recesses which extend perpendicularly with respect to the surface in a turbine blade or vane and have an outer thermal barrier coating on a bonding layer.

SUMMARY OF THE INVENTION

Therefore, it is an object of the invention to provide a component having improved thermal insulation and to specify a process for operating a turbine which reduces the need to cool components.

The object is achieved by means of a component having recesses extending diagonally with respect to the direction of flow and by means of a process for operating a turbine comprising such components.

The recesses preferably extend only in one layer, i.e. are preferably present within one layer.

If a plurality of, in this case preferably two, layers are present, the recesses are then present only in the outermost layer. In the example of turbine blades or vanes for gas turbines, the outermost layer is a ceramic layer in which the recesses are present.

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

BRIEF DESCRIPTION OF THE DRAWINGS

The invention is to be described in more detail below with reference to appended drawings.

In the figures:

FIGS. 1, 2, 3, 4, 5, 6 show exemplary embodiments,

FIG. 7 shows a gas turbine,

FIG. 8 shows a perspective view of a turbine blade or vane, and

FIG. 9 shows a perspective view of a combustion chamber.

DETAILED DESCRIPTION OF THE INVENTION

FIG. 1 shows a component 1 in cross section.

Particularly in the case of gas turbines 100 (FIG. 7), the component 1 is a turbine rotor blade or guide vane 120, 130 (FIGS. 7, 8) or a combustion chamber element 155 (FIG. 9).

The invention is explained merely by way of example with reference to turbine blades or vanes 120, 130 of gas turbines 100, but may be used for any desired component which has a medium flowing over or around it, that is to say also in gas turbines for aircraft or in steam turbines or compressors.

In particular, the component 1, 120, 130, 155 comprises a substrate 4 which, particularly in the case of high-temperature applications, such as in gas turbines, consists of a nickel-base or cobalt-base alloy. Iron-base superalloys are also used in the case of components of steam turbines.

A bonding layer 7 which preferably consists of an alloy of the MCrAlX type and to which an outer ceramic thermal barrier coating 10 has been applied is preferably present on the substrate 4.

The recesses 19 start from a surface 16 of the component 120, 130, 155 and may be present in a solid component 120, 130, 155 (component comprising only a substrate 4) or in layers 7, 10 (FIGS. 1, 2, 3). The recesses 19 may also extend through one or more layers 7, 10 (not shown).

The component 1 has a medium flowing over or around it in the direction of flow 13. The recesses 19 preferably extend diagonally in the direction of flow 13 (FIGS. 1, 2, 3). Equally, however, they may also extend diagonally counter to the direction of flow 13.

The recess 19 represents a blind hole or always has a base 28. It is therefore not used as a film-cooling hole.

The recesses 19 have a longitudinal direction 22 which extends within the recess 19 from the base 28 of the recess 19 as far as the surface 16 of the component and which extends at an angle α diagonally with respect to the direction of flow 13 or with respect to the surface 16 (FIG. 2).

The penetration depth d of a recess 19 extends perpendicularly with respect to the surface 16 of the component 120, 130 and may be dimensioned in values relative to the layer thickness s of the individual layers 7, 10 and to the overall layer thickness.

A penetration depth d of the recess into one layer 10 or into the layers 7, 10 is preferably defined in values relative to the layer thickness s of the outermost layer. The penetration depth extends perpendicularly with respect to the outer surface 16. It is preferably 10%-120% of the layer thickness s, i.e. in the case of 120%, it extends into the substrate 4 or an underlying and/or underlying layer 7 and into the substrate 4 via the outer layer 10.

The penetration depth d is preferably between 10% and 90% of the layer thickness s of the outermost layer 10, i.e. it is arranged only within the outermost layer 10. Particular preference is given to using penetration depths of 50%-80% of the layer thickness of the outermost layer 10 (FIG. 3).

The outermost layer 10 preferably has a thickness of from 1-2 mm and, for the recess 19, has a penetration depth d of 1 mm.

The recesses 19 preferably have the same penetration depth d (FIG. 2) from the surface 16 of the component. A penetration depth d is preferably from 10% to 120% of the layer thickness s.

The angle α is not 90° (α≠90°, i.e. α>90° or α<90°). The difference from 90° is selected such that it is outside a tolerance range given for the production of perpendicularly extending recesses, as is known from U.S. Pat. No. 6,703,137 B2.

The angle α is preferably <80° or >100°.

The angle α is preferably between 20° and 80°.

The recess 19 is preferably of elongated form in the plane of the surface 16 of the component 1, 120, 130, 155, i.e. the extent 1 in the plane of the surface 16 is preferably at least ten times greater than the penetration depth d (FIGS. 3, 4).

The recess 19 may also be bent (FIG. 4).

The recess 19 may also surround a component 120, 130, 155, i.e. may surround the main blade or vane part 406 in the case of a turbine blade or vane 120, 130.

The recess 19 may have a medium flowing over it at an angle of β=90°±90° (FIGS. 3, 5): 0°<β<180°, in particular 10°<β<170°. The angle β is defined by the direction of flow 13 and a lateral direction 25, which represents an edge of the recess 19 level with the surface 16.

FIG. 2 shows a further exemplary embodiment.

Starting from FIG. 1, film-cooling holes 418, 419 are present in the substrate 4 and/or also in the layers 7, 10. The film-cooling holes 418 extend from a cavity of the component preferably until they are level with the penetration depth d of the recesses 19.

The film-cooling holes 418 may also extend as far as the surface 16 (not shown), where recesses 19 are located or else where no recesses 19 are located.

It is also possible for concealed film-cooling holes 419 to be present, and these are present underneath the thermal barrier coating 10 and underneath the bonding layer 7.

The film-cooling hole 418 may be as wide as the recess at the level of the plane 20, and may be thinner or else wider than the extent of the recess 19 in the direction of flow 13.

The recess may have any desired cross section.

In FIGS. 1, 2, the recesses are in the form of a parallelogram. In cross section parallel to the surface 16, the edges of the recess 19 have edges extending in parallel in cross section.

The recess 19 may also be wider in the region of the surface 16 than in the region of the base 28 of the recess 19 (FIG. 4).

The width of the recess 19″ on the surface 16 may also be smaller than on the base 28 level with the penetration depth d.

The longitudinal direction 22 is always formed by a line which extends in the plane of a side wall 23, 26 and has the smallest distance between the base 28 of the recess 19 and the surface 16 of the recess 19.

The recesses 19 may be introduced in different ways. In the case of metallic layers 7 or metallic substrates 4, this can be done using a known mechanical method. In the case of ceramics and under ceramic layers 10, this is preferably done by means of a laser, as is also explained in U.S. Pat. No. 6,703,137 B2, or by means of electron irradiation.

The recesses 19 have the effect that the air molecules do not move and thus form a type of open porosity, in which case the air remains in the recesses or slots 19 as a result of the diagonal position in the direction of flow 13.

FIG. 6 shows by way of example a partial longitudinal section through a gas turbine 100.

In its interior, the gas turbine 100 has a rotor 103 which is mounted such that it can rotate about an axis of rotation 102, has a shaft 101, and is also referred to as the turbine rotor.

An intake casing 104, a compressor 105, a for example toric combustion chamber 110, in particular an annular combustion chamber, with a plurality of coaxially arranged burners 107, a turbine 108 and the exhaust gas casing 109 follow one another along the rotor 103.

The annular combustion chamber 110 is in communication with a for example annular hot gas duct 111. There, by way of example, four successive turbine stages 112 form the turbine 108.

Each turbine stage 112 is formed for example from two blade rings. As seen in the direction of flow of a working medium 113, a guide vane row 115 is followed in the hot gas duct 111 by a row 125 formed from rotor blades 120.

The guide vanes 130 are secured to an inner casing 138 of a stator 143, whereas the rotor blades 120 belonging to a row 125 are arranged on 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, air 135 is drawn in through the intake casing 104 and compressed by the compressor 105. The compressed air provided at the turbine end of the compressor 105 is passed to the burners 107, where it is mixed with a fuel. The mixture is then burnt in the combustion chamber 110, forming the working medium 113. From there, the working medium 113 flows along the hot gas duct 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 can be cooled by means of a coolant.

Substrates of the components may likewise have a directional structure, i.e. they are in single-crystal form (SX structure) or have only longitudinally oriented grains (DS structure).

By way of example, iron-base, nickel-base or cobalt-base 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; these documents form part of the disclosure with regard to the chemical composition of the alloys.

The guide vane 130 has a guide vane root (not shown here) facing the inner casing 138 of the turbine 108 and a guide vane head 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.

FIG. 7 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 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, 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.

A blade or vane root 183, which is used to secure the rotor blades 120, 130 to a shaft or a disk (not shown), is formed in the securing region 400.

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; these documents form part of the disclosure with regard to the chemical composition of the alloy.

The blade or vane 120, 130 may in this case be produced by a casting process, also 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; these documents form part of the disclosure with regard to the solidification process.

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 of the rare earth elements, or hafnium (Hf)). Alloys of this type are known from EP 0 486 489 B1, EP0 786 017 B1, EP0 412 397 B1 or EP 1 306 454 A1, which are intended to form part of this disclosure with regard to the chemical composition of the alloy.

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 interlayer or as the outermost layer).

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 means of suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD).

Other coating processes are conceivable, for example atmospheric plasma spraying (APS), LPPS, VPS or CVD. The thermal barrier coating may have grains that are porous and/or include micro-cracks or macro-cracks in order to improve the resistance to thermal shocks. Therefore, the thermal barrier coating is preferably more porous than the MCrAlX layer.

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).

FIG. 8 shows a combustion chamber 110 of the gas turbine 100. The combustion chamber 110 is configured, for example, as what is known as an annular combustion chamber, in which a multiplicity of burners 107, which generate flames 156 and are arranged circumferentially around an axis of rotation 102, open out into a common combustion chamber space 154. For this purpose, the combustion chamber 110 overall is of annular configuration positioned around the axis of rotation 102.

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.

A cooling system may also 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) which open out into the combustion chamber space 154.

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 those used for 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 of the rare earth elements, 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, which are intended to form part of this disclosure with regard to the chemical composition of the alloy.

A for example ceramic thermal barrier coating, 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, may also be present on the MCrAlX.

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 conceivable, for example atmospheric plasma spraying (APS), LPPS, VPS or CVD. The thermal barrier coating may have grains that are porous and/or include 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 turbine blades or vanes 120, 130, 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 turbine blade or vane 120, 130 or the heat shield element 155 are also repaired. This is followed by recoating of the turbine blades or vanes 120, 130, heat shield elements 155, after which the turbine blades or vanes 120, 130 or the heat shield elements 155 can be reused.

Claims

1.-16. (canceled)

17. A component of a turbomachine having a medium flowing over the component in a direction of flow, comprising:

an outer surface; and

a recess presented within a layer on the component that extends through a longitudinal direction from a base of the recess to the outer surface at an angle not perpendicular with respect to the outer surface.

18. The component as claimed in claim 17, wherein the longitudinal direction extends in a plane of a side wall of the component.

19. The component as claimed in claim 17, wherein the angle is less than 85° or greater than 95°.

20. The component as claimed in claim 17, wherein the recess is elongated in a plane of the outer surface transverse with respect to the direction of flow.

21. The component as claimed in claim 17, wherein the recess is present in a ceramic layer on the component.

22. The component as claimed in claim 17, wherein the recess is present in an outermost layer on the component.

23. The component as claimed in claim 22, wherein a penetration depth of the recess perpendicular with respect to the outer surface is from 10% to 120% of a layer thickness of the outermost layer.

24. The component as claimed in claim 23, wherein the penetration depth is from 10% to 90% of the layer thickness of the outermost layer.

25. The component as claimed in claim 23, wherein the penetration depth is from 50% to 80% of the layer thickness of the outermost layer.

26. The component as claimed in claim 17, wherein the recess is a parallelogram and has edges extending in parallel to the outer surface in a cross section.

27. The component as claimed in claim 17, wherein the component has cooling holes in a region of the outer surface with the recess.

28. The component as claimed in claim 17, wherein the component is a turbine component.

29. The component as claimed in claim 28, wherein the component is a turbine guide vane or rotor blade of a gas turbine.

30. A component of a turbomachine having a medium flowing over the component in a direction of flow, comprising:

an outer surface; and

a recess presented within a layer on the component that extends through a longitudinal direction from a base of the recess to the outer surface at an angle not perpendicular with respect to the direction of flow.

31. The component as claimed in claim 30, wherein the angle is less than 85° or greater than 95°.

32. The component as claimed in claim 31, wherein the angle is less than 80°.

33. The component as claimed in claim 31, wherein the angle is greater than 100°.

34. A method for operating a turbine, comprising:

flowing a medium over a component of the turbine in a direction of flow;

presenting a recess within a layer on the component; and

extending the recess through a longitudinal direction from a base of the recess to an outer surface of the component at an angle not perpendicular with respect to the outer surface.