Abstract:
The invention relates to a turbine blade ( 31 ) having a root region ( 33 ), a tip region ( 37 ) and a blade airfoil ( 35 ), a rounded portion ( 73 ) being formed between the root region ( 33 ) and the blade airfoil ( 35 ), and relief slots ( 51 ) passing through the blade trailing edge ( 41 ) in the region of this rounded portion ( 73 ), thermal expansions being compensated for and thus thermal stresses being minimized by these relief slots ( 51 ).

Description:
CROSS REFERENCE TO RELATED APPLICATION 
   This application claims priority of the European application No. 03020211.3 EP filed Sep. 5, 2003, which is incorporated by reference herein in its entirety. 
   FIELD OF THE INVENTION 
   The invention relates to the blade of a turbine, which blade is directed along a blade axis, is formed from a basic body and comprises a root region, a tip region and a blade airfoil having an airfoil height extending from the root region to the tip region. The invention also relates to a method of preventing the propagation of cracks in the blade airfoil of the blade of a turbine. 
   BACKGROUND OF THE INVENTION 
   In turbines, a flow medium is transported in a flow duct in order to obtain energy therefrom. To this end, turbine blades are arranged in the flow duct. For example, in the flow duct of an axial gas turbine, guide-blade rings formed from guide blades and moving-blade rings formed from moving blades are arranged alternately following one another in the direction of flow. In a suitable manner, the guide blades deflect the flow medium onto the moving blades, which are connected to a rotor and are set in rotation, so that kinetic energy of the flow medium is converted into rotational energy. 
   Such blades in fluid-flow machines are often subjected to considerable mechanical loads. Especially at a simultaneous high temperature and high rotary speeds, as in a gas turbine, the blade material is subjected to high stress. As a result, cracks may form in the blade material, and these cracks propagate in the course of time when stress is continuous. Finally, failure of the blade may occur, the blade breaking into pieces or fragments being released. For blades following in the direction of flow, this may lead to considerable damage. The formation and propagation of cracks thus need to be monitored. Depending on the speed of the processes, a significant reduction in the availability of the turbine may occur as a result, since regular service intervals lead to turbine downtimes. 
   Described in U.S. Pat. No. 6,490,791 is a method in which cracks in the trailing edge of a turbine blade are removed in a service process by cutting back the trailing edge. The additional aerodynamic losses caused by the shortened trailing edge are kept low by subsequent rounding of the blade profile. Although this method can avoid a complete exchange of used blades for new blades, it does not reduce the frequency of the service intervals. 
   Shown in JP 2000018001 is a gas-turbine moving blade in which relief slots are incorporated in the direction of the blade axis toward the margin of the tip region. These relief slots serve to reduce thermal stresses in this region. The reduction in thermal stresses is intended to reduce crack formation. The relief slots are restricted to the tip region. 
   JP 10299408 shows a gas-turbine blade in which elliptical holes which are intended to reduce crack propagation are incorporated in regions of high thermal stresses. The holes are arranged in the transition region between blade airfoil and platform, the major axis of the ellipse being directed perpendicularly to the blade axis in the region of the blade airfoil. There is a corresponding orientation of the holes at the trailing edge. 
   SUMMARY OF THE INVENTION 
   The object of the invention is to specify a turbine blade which is subjected to especially low thermal stresses. 
   According to the invention, this object is achieved by a turbine blade which is directed along a blade axis, is formed from a basic body and comprises a root region, a tip region and a blade airfoil having an airfoil height extending from the root region to the tip region and an airfoil width extending from a blade leading edge up to a blade trailing edge, a rounded portion being formed in a transition region between the blade trailing edge and the root region, a relief slot being formed transversely through the blade trailing edge. 
   The invention is based on the knowledge that the blade trailing edge of a turbine blade is itself subjected to especially high mechanical stresses in a region above the rounded transition region between the blade trailing edge and the root region and in this rounded transition region itself. Furthermore, the invention is based on the knowledge that, given appropriate dimensioning, the blade trailing edge is not unduly destabilized mechanically by slots which run transversely to it. By the introduction of a relief slot transversely to and through the blade trailing edge, considerable relief from thermal stresses is now achieved owing to the fact that thermal expansion can be compensated for by the slot. 
   The relief slot preferably lies in the vicinity of the rounded portion. The blade trailing edge is subjected to especially high thermal stresses especially in a region in the vicinity of the rounded portion. The stresses in this especially affected region can be effectively reduced by the relief slot. Furthermore, in this case, the relief slot is preferably at a distance from the rounded portion of less than 20% of the airfoil height. A distance of the relief slot from the rounded portion of less than 10% of the airfoil height is especially preferred. 
   The slot preferably has a length of at least 2% of the airfoil width. With a slot of this extent, an especially high relief effect is achieved by means of the slot. 
   The slot preferably has at most a length of 5% of the airfoil width. A slot length having an extent greater than 5% of the slot width only leads to a comparatively small further relief of thermal stresses, whereas on the other hand the mechanical stability of the blade trailing edge would suffer by an excessive slot length. 
   Preferably at least two, more preferably at least three, relief slots are provided. With more than two or three relief slots following one another along the blade axis, a larger region of the blade trailing edge can be relieved of thermal stresses. In addition, higher thermal stresses overall can be countered. All the relief slots are preferably at a distance from the rounded portion within a range of less than 25% of the airfoil height. 
   Three relief slots are preferably provided, a first slot nearest to the rounded portion having a first length, a second slot following the first slot along the blade axis having a second length, and a third slot following the second slot along the blade axis having a third length, the third length being greater than the second length, and the second length being greater than the first length. 
   The turbine blade is preferably a gas-turbine blade. Gas turbines are exposed to especially high temperatures. Accordingly, the build-up of especially high thermal stresses occurs precisely in this case. 
   The relief slot preferably has an approximately circular widened portion at its end opposite the blade trailing edge. Due to such a circular widened portion, the radii of curvature of the surfaces defining the slot at the end are reduced and thus the stresses occurring in particular at such curvatures are reduced. In particular, the circular widened portion is a circular hole, from which the slot extends through the blade trailing edge. The slot is preferably cut by means of a laser beam or it is milled. 
   The measures described above for the relief of thermal stresses in the blade trailing edge are also equally suitable for reducing stresses in the rounded portion between the blade trailing edge and the root region. 
   Very high stresses may occur especially in this rounded portion or also in a notch. This zone is thus also a preferred zone for crack formation. 

   
     BRIEF DESCRIPTION OF THE DRAWINGS  
     The invention is explained in more detail by way of example with reference to the drawings. 
     In the drawings, partly schematically and not to scale: 
       FIG. 1  shows a gas turbine, 
       FIG. 2  shows a gas-turbine guide blade, and 
       FIG. 3  shows a detail of a longitudinal section through a gas-turbine guide blade in the region of the rounded portion between blade trailing edge and root region. 
   

   DETAILED DESCRIPTION OF INVENTION 
   The same reference numerals have the same meaning in the various figures. 
     FIG. 1  shows a gas turbine  1 . The gas turbine  1  is directed along a turbine axis  10  and has, following one another along the turbine axis  10 , a compressor  3 , a combustion chamber  5  and a turbine part  7 . The compressor  3  and the turbine part  7  are arranged on a common turbine shaft  9 . Formed in the turbine part  7  is a hot-gas duct  12 , into which guide blades  11  and moving blades  13 , which are arranged on the turbine shaft  9 , project. 
   During operation of the gas turbine  1 , ambient air is drawn in by the compressor  3  and compressed to form compressor air  15 . The compressor air  15  is burned with fuel in the combustion chamber  5  to form hot gas  17 , which flows through the hot-gas duct  12 . In the process, the turbine shaft  9  is set in motion via the effect on the moving blades  13 . The rotational energy of the turbine shaft  9  can be used, for example, for generating electrical energy. 
     FIG. 2  shows a gas-turbine guide blade  31 . The gas-turbine guide blade  31  has a root region  33  with a platform  34 . A blade airfoil  35  adjoins the platform  34 . The blade airfoil  35  ends in a tip region  37 , which in particular also has a platform, which, however, is not shown here. The platform  34  and also the platform (not shown) of the tip region  37  serve to define the hot-gas duct  12 . The blade airfoil  35  has an airfoil height h. Furthermore, the blade airfoil  35  has a blade width b. The blade airfoil  35  extends from a blade leading edge  39  to a blade trailing edge  41 . The pressure side  45 , on the one hand, and the opposite suction side  47 , on the other hand, of the blade airfoil  35  lie between blade leading edge  39  and blade trailing edge  41 . The gas-turbine guide blade  31  has a basic body  32  which is of hollow design, a blade outer wall  63  enclosing the cavity. Stabilizing side ribs  65  are arranged in the cavity between the suction side  47  and the pressure side  45 . 
   In the transition region between blade airfoil  35  and platform  34 , a rounded portion  71  is formed in the region of the blade leading edge  39  and a rounded portion  73  is formed in the region of the blade trailing edge  41 . These rounded portions  71 ,  73 , also referred to as thickened portions or notches, are subjected to especially high mechanical stresses during operation. For relief from thermal stresses which occur due to the high temperatures to which the gas-turbine guide blade  31  is exposed, relief slots  51  are provided in the blade trailing edge. These relief slots  51  are described in more detail with reference to  FIG. 3 . 
     FIG. 3  shows a detail of a longitudinal section through the gas-turbine guide blade  31  in the region of the rounded portion  73  between blade trailing edge  41  and platform  34 . The relief slots  51  extend transversely to and through the blade trailing edge  41 . In this case, the blade trailing edge  41  may be formed, for example, solely by the suction side  47 , whereas the pressure side  45  ends in a stepped manner, and cooling-air openings which cool the blade trailing edge  41  are provided in this step. This would be an open blade trailing edge  41 . However, there may also be a closed blade trailing edge  41 , in which the pressure side  45  merges in a rounded manner into the suction side  47  and forms the blade trailing edge  41  in the process. In this case, the relief slots  51  may extend in the suction side  47 , the pressure side  45  or in both sides. With their ends opposite the blade trailing edge  41 , the relief slots  51  end in circular widened portions  53 , in which comparatively few stresses are caused due to a relatively small curvature. The relief slot  51  nearest to the rounded portion has a smaller volume than the second relief slot following in the blade axis direction. The second relief slot is in turn shorter than the third relief slot  51  which follows it in the direction of the blade axis and is furthest away from the rounded portion  73 . 
   Thermal stresses are reduced by the relief slots  51  by virtue of the fact that a thermal expansion can be compensated for in the relief slots  51 . As a result, thermal stresses both in the region of the trailing edge  41  and in the rounded portion  73  are minimized. 
   Cooling air  67  for the cooling is directed into the gas-turbine guide blade  34 . This cooling air  67  comes out of the slot  51  from the hollow interior of the gas-turbine guide blade  34 . In this case, the slot  51  is shaped in such a way that the cooling air  67  forms a cooling film on the surface of the blade airfoil  35 .