Abstract:
An aspect of the invention is a turbine blade for a gas turbine, comprising a blade root, adjoining which one after the other are a platform region having a transversely running platform and then a blade profile curved in the longitudinal direction, comprising at least one cavity which is open on the root side and through which a coolant can flow and which extends through the blade root and the platform region into the blade profile. The cavity is surrounded by an inner wall, on the surface of which structural elements influencing the coolant are provided. In order to prolong the service life of such a turbine blade, the invention proposes that a section, lying at least in the blade profile and adjoining the platform region, of the surface of the inner wall be free of structural elements. Such a turbine blade can preferably be used in a stationary gas turbine.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     This application is the US National Stage of International Application No. PCT/EP2006/064414, filed Jul. 19, 2006 and claims the benefit thereof. The International Application claims the benefits of European application No. 05016328.6 filed Jul. 27, 2005, both of the applications are incorporated by reference herein in their entirety. 
     FIELD OF INVENTION 
     The invention relates to a turbine blade for a gas turbine, with a blade root, to which a platform region, with a transversely extending platform, and upon it a blade airfoil, which is curved in the longitudinal direction, are connected in succession, with at least one cavity which is open on the root side, is exposable to throughflow by a cooling medium, and extends through the blade root and the platform region into the blade airfoil. Furthermore, the invention relates to the use of such a turbine blade. 
     BACKGROUND OF THE INVENTION 
     A cooled rotor blade of a gas turbine, which inside has cooling passages which extend in meander-form, is known from EP 1 469 163 A2. Turbulators, which stimulate the heat transfer from blade material to the cooling medium which flows through the cavity, are provided on the inner walls which delimit the cavities, in the region of the blade airfoil. As a result of the increased heat transfer, the turbine blade can consequently withstand higher operating temperatures. 
     In this case, it is disadvantageous that cracks can occur in the region of the fillet-like transition from platform to the blade airfoil, which transition in English is also referred to as a fillet, and/or in the platform. If the cracks which develop exceed a critical crack length, then a safe operation of the gas turbine, which is equipped with such a turbine blade, is not ensured. 
     SUMMARY OF INVENTION 
     Therefore, an especially long service life of the turbine blade is a design objective, by which the availability duration of a gas turbine which is equipped with it can be further increased. The object of the invention is the provision of a turbine blade for a gas turbine, with which the fatigue life is extended. Moreover, it is the object of the invention to disclose the use of such a turbine blade. 
     The object which is focused upon the turbine blade is achieved by a generic-type turbine blade which is designed according to the features of the claims. 
     The invention is based on the knowledge that wear and crack development, and also the subsequent crack propagation, are thermally dependent. The material of the turbine blade is subjected to thermal stresses which arise as a result of the external impingement by hot gas and the cooling which takes place inside. It has been proved that during operation of the gas turbine, locally comparatively low temperatures on the hot gas side occur in the fillet-like transition region between blade airfoil and platform, compared with those temperatures in the region of the blade airfoil. Therefore, the internally cooled turbine blade, with turbulators which are arranged on the inner walls in the region of the platform, was previously cooled too intensely in locally defined regions. Consequently, locally comparatively large temperature differences and correspondingly large thermal stresses, which were able to cause wear, occurred in the blade material. 
     The invention proposes to significantly reduce these local thermal stresses in the transition region by this simply not being cooled as intensely as the blade airfoil. In order to achieve this, with a generic-type turbine blade it is provided that a section of the surface of the inner wall, which section lies at least within the blade airfoil and adjoins the platform region, is free of structural elements. 
     Consequently, the heat transfer from the blade material to the cooling medium which flows past is locally reduced in the region of the transition radius in order to purposefully reduce in this way the thermal gradient at this point. This reduction leads to a locally hotter transition region with regard to the prior art. Therefore, lower thermal stresses are formed in the transition radius between the platform and the blade airfoil, as a result of which at this point crack development can be reduced and crack propagation can be delayed. Wear is reduced as a result. 
     At the same time, the temperature drop in the blade material, on account of the hotter transition region, is lowered in the section between the edge of the platform and the cavity, which extends the service life of the turbine blade. 
     By means of the proposed measure, the service life, especially the fatigue life (low cycle fatigue=LCF) for the platform and its transition into the blade airfoil, i.e. in the fillet, is extended. 
     Advantageous developments are disclosed in the dependent claims. 
     The development in which the surface of the inner wall at the level of the platform region, and the surface of the inner wall of the section which adjoins it inside the blade airfoil, are flat, is especially advantageous. On account of the unswirled cooling medium flow in this section, the heat transfer from the blade material to the cooling medium, compared with the heat transfer in the blade airfoil, is reduced, so that the temperature difference between an external surface of the blade airfoil, which is impinged by hot gas, that is the hot side, and the inner wall of the turbine blade which is impinged by cooling medium, that is the cold side, can be significantly reduced by means of a permissible raising of the material temperature. The reduction leads to reduced thermal stresses, especially in the region of the transition between the blade airfoil and the platform, that is in the fillet. 
     Since the structural elements on the inner wall of the blade airfoil as a rule are indeed areally spaced apart, but, as viewed in the radial direction, forming a mean minimum spacing, an advantageous development provides that a distance which is defined between the platform surface and, also as viewed in the radial direction, the adjacent structural element nearest to it, is greater than the mean minimum spacing between two adjacent structural elements. In this case, the distance is preferably at least 1.1 times the mean minimum spacing. 
     It has been proved to be further advantageous for the section to have a height of 5% of the airfoil height of the blade airfoil up to the airfoil tip, calculated from the platform surface. The development in which a region of the inner wall, which has the structural elements and lies within the blade airfoil, starts only after a height of 10% of the airfoil height, calculated from the platform surface in the direction of the airfoil tip, is especially advantageous. 
     An especially advantageous reduction of the temperature difference between the hot side and the cold side, especially in the otherwise especially wear-affected transition region, can be effected by means of these measures. 
     In an advantageous development, the structural elements are formed as turbulators in the form of ribs, block fields, dimples and/or nipples. 
     Since the local temperature difference between the hot side and the cold side, which causes wear, especially in a center region of the transition region, occurs between a leading edge of the blade airfoil and a trailing edge of the blade airfoil, it is especially advantageous if the surface of the inner wall which lies in the center region between the leading edge and the trailing edge is free of structural elements. In this case, the turbine blade can have a plurality of cavities which extend through the turbine blade in the radial direction and are separated by means of support ribs, in which only the cavity which lies between the leading edge and the trailing edge of the blade airfoil in the center region has the section of the inner wall, the surface of the inner wall of which within the blade airfoil is free of structural elements. 
     This comes from the knowledge that along the longitudinal edge of the platform, as viewed from the leading edge to the trailing edge, a temperature variation is established in the blade material, which in the region of the leading edge and of the trailing edge has in each case a relative maximum and between them, in the center region, has a local minimum. This temperature minimum can be raised by means of the proposed measures. As a result, only the regions in which especially high temperature gradients previously occurred, i.e. temperature differences between the hot side and the cold side on account of an excessive cooling, are purposefully locally cooled less. In contrast, the cavities in the region of the leading edge and in the region of the trailing edge which extend along them can be provided as before with structural elements which reach to the platform. 
     The platform which is arranged on the pressure side in the center region between the leading edge and trailing edge is especially wide for structural reasons, so that the local temperature minimum in the blade material previously occurred at this point. The temperature minimum can be raised by reducing the thermal stress if especially the surface of the inner wall, which inner wall is formed by the suction-side airfoil wall of the blade airfoil, is free of structural elements. Consequently, an especially long service life extension of the expediently cast turbine blade can be brought about. 
     Moreover, for achieving the second-mentioned object, the use of a turbine blade as claimed in the claims in a preferably stationary gas turbine is proposed. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The invention is explained with reference to Figures. In the drawing: 
         FIG. 1  shows a gas turbine in a longitudinal partial section, 
         FIG. 2  shows a turbine blade in perspective view with overhanging platform regions, 
         FIG. 3  shows the turbine blade according to the invention in cross section with different cooling configurations, and 
         FIG. 4  shows a turbine blade according to the invention in longitudinal section with turbulators which start at different radial heights. 
     
    
    
     DETAILED DESCRIPTION OF INVENTION 
       FIG. 1  shows a gas turbine  1  in a longitudinal partial section. Inside, it has a rotor  3 , which is also referred to as a turbine rotor, and which is rotatably mounted around a rotational axis  2 . An inlet duct  4 , a compressor  5 , a toroidal annular combustion chamber  6  with a plurality of burners  7  which are arranged rotationally symmetrically to each other, a turbine unit  8  and an exhaust duct  9 , follow in succession along the rotor  3 . The annular combustion chamber  6  forms a combustion space  17  which communicates with an annular hot gas passage  18 . Four turbine stages  10 , which are connected one behind the other, form the turbine unit  8  there. Each turbine stage  10  is formed from two blade rings. In the hot gas passage  18 , a row  14  which is formed from rotor blades  15  follows a stator blade row  13  in each case, as seen in the flow direction of a hot gas  11  which is produced in the annular combustion chamber  6 . The stator blades  12  are fastened on the stator, whereas the rotor blades  15  of a row  14  are attached on the rotor  3  by means of a turbine disc  19 . A generator or a driven machine (not shown) is coupled to the rotor  3 . 
       FIG. 2  shows a hollow turbine blade  50  according to the invention in perspective view. The preferably cast turbine blade  50  comprises a blade root  52  upon which a platform  54 , and upon it a blade airfoil  56 , which is not shown in its full height but shown in a shortened form, are arranged along a blade axis. 
     The blade airfoil  56  has a pressure-side airfoil wall  62 , and also a suction-side airfoil wall  64 , which extend from a leading edge  66  of the blade airfoil  56  to a trailing edge  68 . 
     During operation of the gas turbine  1 , the hot gas  11  flows along the airfoil walls  62 ,  64  from the leading edge  66  in the direction of the trailing edge  68 . 
     A fillet-like transition region  48  is formed between the platform  54  and the blade profile  56 . 
     Three sub-cavities  58 , in which a cooling medium K, which is provided for cooling, can flow in each case, extend through the turbine blade  50  from the blade root  52  into the blade airfoil  56 . The first sub-cavity  58   a  extends parallel to, and in the region of, the leading edge. A second sub-cavity  58   b  follows behind it, as seen in the flow direction of the hot gas. 
     The sub-cavities  58  extend in the radial direction with regard to the installed position of the turbine blade  50  in the gas turbine  1 , and are separated from each other by means of support ribs  70 . For stiffening the blade airfoil  56 , the support ribs  70  connect the pressure-side airfoil wall  62  to the suction-side airfoil wall  64 . 
     On account of the platform longitudinal edge  63 , which is rectilinear in the axial direction, of the rectilinear blade root  52  and of the blade airfoil  56  which is curved in the same direction, the platform surface  61  on the pressure side, in the region of the center sub-cavity  58 , has a width B which extends transversely to the axial direction and is greater than the width of the platform surface  61  which is provided in the pressure-side region of the leading edge  66  or trailing edge  68 . 
     For reasons of clarity, no structural elements are shown in the sub-cavities  58  of the turbine blade  50  which is shown in  FIG. 2 . 
       FIG. 3  shows the turbine blade  50  according to the invention, which is formed as a rotor blade or stator blade, in accordance with the cross section III-III of  FIG. 2 . The platform  54  and the blade airfoil  56  follow the blade root  52  in the radial direction, with regard to the installed position in the gas turbine  1 . Both the outer side of the blade airfoil  56  and the surface  61  of the platform  54  which faces the blade airfoil  56  are subjected to the hot gas  11  which flows through the gas turbine  1 , and are referred to as the hot side. 
     The cutting plane of the cross section III-III extends through the second of the three sub-cavities  58  which in each case are open on the root side. The cooling medium K, for example cooling air, which can be fed on the root side, cools the turbine blade  50  so that this can withstand the temperatures which occur during operation of the gas turbine. 
     The second sub-cavity  58   b  is enclosed by an inner wall  59  which is partially formed by the pressure-side airfoil wall  62  and by the suction-side airfoil wall  64 . For increasing the heat transfer of the blade material, which is heated by the hot gas  11 , to the cooling medium K which flows inside, structural elements  72  in the form of turbulators, which can be formed as ribs, block fields, dimples and/or nipples, are provided on the inner surfaces of the airfoil walls  62 ,  64  or of the inner walls  59 . In the development which is shown, they are ribs which extend transversely to the direction of cooling medium flow. 
     Previously, it was customary to provide the turbulators or the structural elements  72  approximately over an entire airfoil height H from the platform  54  to the blade tip  74  ( FIG. 4 ) on the surfaces of the inner walls  59 , just like it is shown on the pressure-side airfoil wall  62  in a first section. A new method is now adopted by the invention. As shown on the inner surface of the suction-side airfoil wall  64 , the structural elements  72  no longer start in the region of the platform surface  61 , but start only after a predetermined height in the blade airfoil  56 . As a result, a second section A of the surface of the suction-side inner wall  59 , which lies within the blade airfoil  56  and adjoins the platform region, is free of structural elements  72 . Although the second section A which adjoins the platform region already lies within the blade airfoil  56 , the surface of the inner wall  59  which is located in this region is correspondingly flat and not profiled by structural elements. 
     A region C of the surface of the inner wall  59 , in which turbulators or structural elements  72  have a mean minimum spacing m in relation to each other, which is defined in the radial direction, adjoins the second section A in the direction of the airfoil tip  74 . 
     On the inner surface of the suction-side airfoil wall  64 , which in the second section A which is close to the platform is free of structural elements  72 , the distance D, which is measured in the radial direction, between the lowermost structural element  73 , or the structural element which is adjacent to the platform surface  61 , and the platform surface  61 , is greater than the mean minimum spacing m. The cooling medium K, which flows in on the root side, first of all flows laminarly in the second section A on account of the locally even base surface and in the meantime convectively cools the blade material. The cooling medium K which flows in the region C is then swirled due to the structural elements  72 ,  73 , which leads to an improved heat transfer. Consequently, it is ensured that the transition region  48  is locally cooled less than the rest of the blade airfoil  56 , and in this way the thermal stresses at this point are reduced, as a result of which cracks only rarely appear. Crack propagation progresses in a delayed manner compared with a turbine blade of the prior art. Consequently, service life of the turbine blade  50  is extended by means of the proposed measures. 
       FIG. 4  shows a further turbine blade  50  according to the invention in longitudinal section, with a blade root  52 , a platform  54  and a blade airfoil  56 . The profiled blade root  52  can be formed in fir-tree form or dovetail form in cross section. The turbine blade  50  is also formed hollow and has four sub-cavities  58  which extend in the radial direction and are separated from each other by means of support ribs  70  which connect the pressure-side airfoil wall  62  to the suction-side airfoil wall  64 . 
     During operation of the gas turbine  1 , a local temperature minimum occurs in the blade material between the front region and the rear region of the transition region  48  on account of the especially wide platform  54  (see  FIG. 2 ) at this point, which blade material is cooled less according to the invention by the structural elements  72  in the two center sub-cavities  58  not starting in the region of the platform surface  61  but starting only from a predetermined height in the blade profile  56 . Therefore, the section A of the surface of the inner walls  59  which are formed by the suction-side airfoil wall  64 , which section lies within the blade airfoil  56  and adjoins the platform region, is free of structural elements  72 . 
     Although the second section A which adjoins the platform region already lies within the blade airfoil  56 , the surface of the inner wall  59  which is located within this region is flat and is not profiled by structural elements. The second section A for example has a height of 5% of the airfoil height H, calculated from the platform surface  61 . The section C of the inner wall, which has the structural elements  72  and lies within the blade airfoil  56 , preferably starts only after a height of 10% of the airfoil height H, calculated from the platform surface  61  in the direction of an airfoil tip  74 . 
     By the invention it is possible to less intensively cool the transition radius or transition region  48  between the blade airfoil  56  and the platform  54 , and especially locally in the center region between leading edge  66  and trailing edge  68 , so that the transition region is subjected to locally smaller temperature differences between the hot side, i.e. outer side of the turbine blade, and the cold side, i.e. inner side of the turbine blade. The smaller temperature differences reduce the thermal stresses in the blade material in the transition region, so that at this point crack development is reduced and crack propagation is delayed, which significantly increases the fatigue life of the turbine blade  50 . 
     A gas turbine which is equipped with such a blade  50  can consequently be operated longer; the turbine blades  50  which are used have to be checked less frequently for defects such as cracks. As a result, the availability of the gas turbine  1  is significantly increased.