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 
       [0001]    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 
       [0002]    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 
       [0003]    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. 
         [0004]    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. 
         [0005]    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 
       [0006]    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. 
         [0007]    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. 
         [0008]    The recesses preferably extend only in one layer, i.e. are preferably present within one layer. 
         [0009]    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. 
         [0010]    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 
         [0011]    The invention is to be described in more detail below with reference to appended drawings. 
           [0012]    In the figures: 
           [0013]      FIGS. 1 ,  2 ,  3 ,  4 ,  5 ,  6  show exemplary embodiments, 
           [0014]      FIG. 7  shows a gas turbine, 
           [0015]      FIG. 8  shows a perspective view of a turbine blade or vane, and 
           [0016]      FIG. 9  shows a perspective view of a combustion chamber. 
       
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
       [0017]      FIG. 1  shows a component  1  in cross section. 
         [0018]    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 ). 
         [0019]    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. 
         [0020]    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. 
         [0021]    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 . 
         [0022]    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). 
         [0023]    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 . 
         [0024]    The recess  19  represents a blind hole or always has a base  28 . It is therefore not used as a film-cooling hole. 
         [0025]    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 ). 
         [0026]    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. 
         [0027]    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 . 
         [0028]    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 ). 
         [0029]    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. 
         [0030]    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. 
         [0031]    The angle α is not 90° (α≠90°, i.e. α&gt;90° or α&lt;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. 
         [0032]    The angle α is preferably &lt;80° or &gt;100°. 
         [0033]    The angle α is preferably between 20° and 80°. 
         [0034]    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 ). 
         [0035]    The recess  19  may also be bent ( FIG. 4 ). 
         [0036]    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 . 
         [0037]    The recess  19  may have a medium flowing over it at an angle of β=90°±90° ( FIGS. 3 ,  5 ): 0°&lt;β&lt;180°, in particular 10°&lt;β&lt;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 . 
         [0038]      FIG. 2  shows a further exemplary embodiment. 
         [0039]    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 . 
         [0040]    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. 
         [0041]    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 . 
         [0042]    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 . 
         [0043]    The recess may have any desired cross section. 
         [0044]    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. 
         [0045]    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 ). 
         [0046]    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. 
         [0047]    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 . 
         [0048]    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. 
         [0049]    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 . 
         [0050]      FIG. 6  shows by way of example a partial longitudinal section through a gas turbine  100 . 
         [0051]    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. 
         [0052]    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 . 
         [0053]    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 . 
         [0054]    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 . 
         [0055]    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 . 
         [0056]    A generator (not shown) is coupled to the rotor  103 . 
         [0057]    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. 
         [0058]    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. 
         [0059]    To be able to withstand the temperatures which prevail there, they can be cooled by means of a coolant. 
         [0060]    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). 
         [0061]    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 . 
         [0062]    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. 
         [0063]    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 . 
         [0064]      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 . 
         [0065]    The turbomachine may be a gas turbine of an aircraft or of a power plant for generating electricity, a steam turbine or a compressor. 
         [0066]    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 . 
         [0067]    As a guide vane  130 , the vane  130  may have a further platform (not shown) at its vane tip  415 . 
         [0068]    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 . 
         [0069]    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. 
         [0070]    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 . 
         [0071]    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 . 
         [0072]    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. 
         [0073]    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. 
         [0074]    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. 
         [0075]    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. 
         [0076]    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. 
         [0077]    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). 
         [0078]    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. 
         [0079]    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. 
         [0080]    The density is preferably 95% of the theoretical density. 
         [0081]    A protective aluminum oxide layer (TGO=thermally grown oxide layer) is formed on the MCrAlX layer (as an interlayer or as the outermost layer). 
         [0082]    It is also possible for a thermal barrier coating, which is preferably the outermost layer and consists for example of ZrO 2 , Y 2 O 3 —ZrO 2 , 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. 
         [0083]    The thermal barrier coating covers the entire MCrAlX layer. 
         [0084]    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). 
         [0085]    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. 
         [0086]    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). 
         [0087]      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 . 
         [0088]    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 . 
         [0089]    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 . 
         [0090]    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). 
         [0091]    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. 
         [0092]    A for example ceramic thermal barrier coating, consisting for example of ZrO 2 , Y 2 O 3 —ZrO 2 , 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. 
         [0093]    Columnar grains are produced in the thermal barrier coating by suitable coating processes, such as for example electron beam physical vapor deposition (EB-PVD). 
         [0094]    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. 
         [0095]    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.