Abstract:
A gas turbine engine having an axial flow compressor, an annular combustion chamber, a turbine, and a diffuser. The diffuser includes a flow-dividing element formed by an inner deflecting flank and an outer deflecting flank that divides a compressed gas flow into two partial flows at a branching point. The two deflecting flanks define: an angle of less than 90° along at least a portion of the deflecting flanks, and an angle between 15° and 90° between the deflecting flank angle bisector and the turbine longitudinal axis. The deflector also includes a main deflecting region arranged upstream of the branching point and directed at an acute angle from the turbine longitudinal axis toward an inner combustion chamber shell.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
     This application is the US National Stage of International Application No. PCT/EP2004/007947, filed Jul. 16, 2004 and claims the benefit thereof. The International Application claims the benefits of European Patent application No. 03018565.6 EP filed Aug. 18, 2003. All of the applications are incorporated by reference herein in their entirety. 
     FIELD OF THE INVENTION 
     The invention relates to a gas turbine having an annular combustion chamber and a diffuser which is arranged upstream of the latter, can be subjected to flow essentially parallel to a turbine longitudinal axis and is at a smaller distance from the latter than the annular combustion chamber and in which a compressed gas can be divided into partial flows at a branching point. 
     BACKGROUND OF THE INVENTION 
     Gas turbines are used in many sectors for driving generators or driven machines. In this case, the energy content of a fuel is used for producing a rotary movement of a turbine shaft. To this end, the fuel is burned in a combustion chamber, in the course of which air compressed by an air compressor is supplied. The working medium which is produced in the combustion chamber by the combustion of the fuel and is under high pressure and high temperature is directed in the process via a turbine unit, where it expands to perform work, arranged downstream of the combustion chamber. 
     In addition to the output which can be achieved, and in addition to a compact type of construction, an especially high efficiency is normally a design aim when designing such gas turbines. In this case, for thermodynamic reasons, an increase in the efficiency can in principle be achieved by an increase in the outlet temperature with which the working medium flows out of the combustion chamber and into the turbine unit. Temperatures of about 1200° C. up to 1300° C. are therefore aimed at and are also achieved for such gas turbines. 
     At such high temperatures of the working medium, however, the components exposed to said working medium are subjected to high thermal loads. In order to nonetheless ensure a comparatively long service life of the relevant components with high reliability, cooling of the relevant components, in particular of moving and guide blades of the turbine unit, is normally provided. Furthermore, provision can be made to cool the combustion chamber with cooling medium, in particular cooling air. 
     DE 195 44 927 A1 discloses a gas turbine which has an air compressor arranged upstream of a combustion chamber and opening into a diffuser. In the diffuser, a partial flow of the compressed air can be branched off from said diffuser and used for cooling structural parts, for example turbine blades of the gas turbine. However, the branching-off of the cooling air from the diffuser is only suitable for branching off a relatively small partial flow from the air flow leaving the air compressor. On the other hand, the main air flow directed through the diffuser is deflected in the direction of the combustion chamber and fed to the latter as combustion air. It is thus likely that components arranged downstream of the diffuser, i.e. relative to the direction of flow of the working medium flowing through the turbine, can only be cooled to a restricted extent. 
     Furthermore, DE 196 39 623 discloses a gas turbine which has a diffuser and in which the cooling air is bled by means of a tube projecting into the outlet of the diffuser. The compressed air used for combustion in an annular combustion chamber is in this case diverted in the direction of the burner by means of a C-shaped plate. Both during the bleeding of the cooling air and during the directing of the burner air, flow losses may occur, which it is necessary to avoid. 
     SUMMARY OF THE INVENTION 
     The object of the invention is to specify a gas turbine which is equipped with an annular combustion chamber and which enables the compressor air to be directed in a fluidically favorable manner for an especially uniform and effective cooling capacity of thermally loaded components. 
     This object is achieved according to the invention by a gas turbine having the features of the claims. In this case, the gas turbine has an annular combustion chamber and an annular diffuser which is arranged downstream of the latter and at least partly between the turbine longitudinal axis and the annular combustion chamber. In the diffuser, which can be subjected to flow essentially parallel to the turbine longitudinal axis, a compressed gas can be divided into a plurality of partial flows. According to the invention, the diffuser has a main deflecting region which is directed at an acute angle pointing away from the turbine longitudinal axis toward the inner wall of the annular combustion chamber. Arranged downstream of the main deflecting region in the direction of the gas, in particular air, flowing through the diffuser is a branching point at which the gas flowing through the diffuser can be divided into partial flows by means of a flow-dividing element. The annular flow-dividing element of wedge-shaped cross section is arranged between the two diverging walls of the diffuser—the inner wall lying radially on the inside and the outer wall lying radially further on the outside. Two deflecting flanks opposite the walls of the diffuser converge at an acute angle and meet at the branching point. There, they enclose an angle bisector which intersects the turbine longitudinal axis at an acute dividing angle greater than 15°. 
     As viewed in the axial direction, the main deflecting region is arranged downstream of the compressor and upstream of the annular combustion chamber, whereas the flow-dividing element is arranged between the annular combustion chamber and the turbine longitudinal axis. For the gas turbine, this geometry permits a compact design which in particular is shortened in the axial direction. Furthermore, the flow losses in the compressed partial flows of cooling medium are reduced. 
     An especially good cooling capacity of components, in particular of the annular combustion chamber, which are at a radial distance from the turbine longitudinal axis is achieved by the gas flow which flows through the diffuser being directed with a component directed toward the annular combustion chamber. The two partial flows divided in the diffuser are preferably then also used for the combustion. 
     In an advantageous development, the outer wall of the diffuser and the outer deflecting flank, opposite said outer wall, of the flow-dividing element run behind the branching point approximately perpendicularly to the turbine longitudinal axis. This ensures low-loss feeding of the outer partial flow to the outer flow transfer space. Short and direct feeding of the partial flow is accordingly achieved. 
     In gas turbines having a combustion chamber not designed as an annular combustion chamber, e.g. in gas turbines having “can combustion chambers”, the supplying of the outer combustion chamber shell is fairly simple. In gas turbines having can combustion chambers, the individual can-shaped combustion chambers are at a distance from one another in the circumferential direction on a ring concentrically enclosing the turbine longitudinal axis. The feeding of the cooling air to the radially outer combustion chamber shells can then be effected between the individual can combustion chambers. 
     Furthermore, low-loss feeding of the inner partial flow to the inner flow transfer space is ensured by the inner wall of the diffuser and the deflecting flank, opposite said inner wall, of the flow-dividing element running approximately parallel to the turbine longitudinal axis. From the compressor outlet up to the flow transfer space, wavelike directing is proposed for the inner partial flow, this wavelike directing, compared with rectilinear directing, achieving an improvement over rectilinear directing with regard to the pressure losses and the flow losses in the partial flow. 
     According to a preferred configuration, the compressed gas, which leaves the diffuser at the branching point, is directed at the latter directly into the flow transfer space, which produces the fluidic connection to the wall cooling space of the annular combustion chamber. The flow transfer space preferably adjoins the combustion chamber wall on the outside, so that additional cooling of the combustion chamber wall is thereby achieved. 
     The annular combustion chamber is preferably of closed coolable design. In this case, combustion air, as cooling medium, is preferably directed through a wall space of the annular combustion chamber in counterflow to the flue gas. The combustion air flowing through the combustion chamber wall is in this case preferably identical at least to a partial flow of the compressed air which has flowed through the diffuser beforehand. The air flowing through the diffuser is preferably fed completely as cooling air to the wall of the annular combustion chamber and further as combustion air to the annular combustion chamber. In this case, the dividing of the air flow at the branching point of the diffuser serves to supply a plurality of parts of the annular combustion chamber, for example an inner shell or an outer shell, uniformly with cooling air. 
     Provided the annular combustion chamber has an essentially flat combustion chamber rear wall, at least in one section, the expression “wall angle” of the annular combustion chamber refers to that angle which the combustion chamber rear wall encloses with the turbine longitudinal axis. Especially uniform all-over cooling of the combustion chamber wall is preferably achieved by the dividing angle of the flow-dividing element deviating from the wall angle of the combustion chamber rear wall by not more than 20°, in particular by not more than 15°. 
     A tube communicating with the bottom sectional passage is preferably provided in order to bleed cooling air for the turbine. As a result, further dividing of the compressor air flow can be effected. If the tube projects into the bottom sectional passage, and its tube opening faces the flow, the turbine cooling air is tapped in an especially favorable manner. 
     The advantage of the invention lies in particular in the fact that air which is compressed in a gas turbine and which serves as cooling air and then as combustion air is fed with a low pressure loss from an air compressor through a compact diffuser to the annular combustion chamber, a flow-dividing element at the outlet of the diffuser producing a uniform admission of cooling air to the annular combustion chamber. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       An exemplary embodiment of the invention is explained in more detail with reference to a drawing, in which: 
         FIG. 1  shows a half section of a gas turbine, and 
         FIG. 2  shows a diffuser and an annular combustion chamber of the gas turbine according to  FIG. 1 , in cross section. 
     
    
    
     Parts corresponding to one another are provided with the same reference numerals in both figures. 
     DETAILED DESCRIPTION OF THE INVENTION 
     The gas turbine  1  according to  FIG. 1  has a compressor  2  for combustion air, an annular combustion chamber  4  and a turbine  6  for driving the compressor  2  and a generator (not shown) or a driven machine. To this end, the turbine  6  and the compressor  2  are arranged on a common turbine shaft  8 , which is also designated as turbine rotor, and to which the generator or the driven machine is also connected, and which is rotatably mounted about its center axis  9 . 
     The annular combustion chamber  4  is fitted with a number of fuel injectors  10  for burning a liquid or gaseous fuel. Furthermore, it is provided with a wall lining  24  at its combustion chamber wall  23 . 
     The turbine  6  has a number of rotatable moving blades  12  connected to the turbine shaft  8 . The moving blades  12  are arranged in a ring shape on the turbine shaft  8  and thus form a number of moving blade rows. Furthermore, the turbine  6  comprises a number of fixed guide blades  14 , which are likewise fastened in a ring shape to an inner casing  16  of the turbine  6  while forming moving blade rows. The moving blades  12  serve in this case to drive the turbine shaft  8  by impulse transmission of the flue, gas or working medium M flowing through the turbine  6 . The guide blades  14 , on the other hand, serve to direct the flow of the working medium M between in each case two successive moving blade rows or moving blade rings as viewed in the direction of flow of the working medium M. A successive pair consisting of a ring of guide blades  14  or a guide blade row and of a ring of moving blades  12  or a moving blade row is designated in this case as a turbine stage. 
     Each guide blade  14  has a platform  18 , which is also designated as blade root  19  and is intended for fixing the respective guide blade  14  in the gas turbine  1 . Each moving blade  12  is fastened to the turbine shaft  8  in a similar manner via a blade root  19  also designated as platform  18 , the blade root  19  in each case carrying a profiled airfoil  20  extended along a blade axis. 
     Between the platforms  18 , arranged at a distance apart, of the guide blades  14  of two adjacent guide blade rows, a respective guide ring  21  is arranged on the inner casing  16  of the turbine  6 . The outer surface of each guide ring  21  is in this case likewise exposed to the hot working medium M flowing through the turbine  6  and is at a radial distance from the outer end  22  of the moving blade  12  lying opposite it with a gap in between. In this case, the guide rings  21  arranged between adjacent guide blade rows serve in particular as cover elements which protect the inner wall  16  or other built-in casing components from thermal overstressing by the hot working medium M flowing through the turbine  6 . 
     To achieve a comparatively high efficiency, the gas turbine  1  is designed for a comparatively high discharge temperature of about 1200° C. to 1300° C. of the working medium M discharging from the annular combustion chamber  4 . 
     The combustion chamber wall  23  can be cooled with cooling air, as cooling medium K, compressed in the compressor  2 . Between the combustion chamber wall  23  and the wall lining  24 , cooling air K flows to the fuel injector  10  in a wall space or wall lining space  26  in counterflow to the working medium M. The cooling air K, which also serves as combustion air, is directed from the compressor  2  through a diffuser  27  in the direction of the annular combustion chamber  4 . By means of the diffuser  27 , the cooling and combustion air K, divided in a defined manner, is fed to an outer combustion chamber shell  28  on the one hand and to an inner combustion chamber shell  29  on the other hand. 
     The directing of the flow of the cooling air K through the diffuser  27  is shown in detail in  FIG. 2 . The diffuser  27  has a main deflecting region  30 , which adjoins the compressor  2 . The compressed cooling air K flows out of the compressor  2  parallel to the center axis or turbine longitudinal axis  9  and into the main deflecting region  30  of the diffuser  27 . The main deflecting region  30 , arranged between the compressor  2  and the annular combustion chamber  4  as viewed in the axial direction, of the diffuser  27  runs radially outward with widening cross section, i.e. away from the turbine longitudinal axis  9 . In this way, the flow velocity of the compressed gas used as cooling air K is reduced in the main deflecting region  30 . Provided a separation of flow occurs at the inner wall and outer wall of the diffuser  27 , such a separation occurs only at a low flow velocity and correspondingly low pressure loss. 
     A flow-dividing element  32  is arranged at the downstream end  31 , with respect to the cooling air K, of the main deflecting region  30  in such a way as to adjoin the outer combustion chamber shell  29 . 
     The flow-dividing element  32  arranged between the annular combustion chamber  4  and the turbine longitudinal axis  9  has an approximately triangular shape, also designated as dividing fork  33 , having an outer deflecting flank  34  and an inner deflecting flank  35 . The deflecting flanks  34 ,  35  converge at a dividing tip  36  directed toward the main deflecting region  30  and enclose an acute angle of less 90°, in particular an angle of 60°, at the dividing tip  36 . The dividing tip or edge  36 , which forms a branching point, divides the cooling air K flowing through the main deflecting region  30  of the diffuser  27  approximately uniformly into an outer cooling air flow K a  and an inner cooling air flow K i . The outer cooling air flow K a  is directed through an outer flow transfer space  37  to an outer combustion chamber shell  28 , whereas the inner cooling air flow K i  is directed via an inner flow transfer space  38  to the inner combustion chamber shell  29 . 
     The diffuser  27  dividing the cooling air K at the flow-dividing element  32  is also designated as split diffuser. The cooling air K flowing through the main deflecting region  30  is deflected radially approximately in a C shape, relative to the turbine longitudinal axis  9 , outward up to the dividing tip  36  of the flow-dividing element  32 . A straight line running as angle bisector  39  between the curved deflecting flanks  34 ,  35  through the dividing tip  36  encloses a dividing angle a of about 45° with the turbine longitudinal axis  9 . The angle bisector  39  encloses an approximately right angle with the bottom combustion chamber shell  29 . The inner cooling air flow K i , starting from the dividing tip  36 , is forced first of all into a horizontal direction of flow, i.e. parallel to the turbine longitudinal axis  9 , by the inner deflecting flank  35  and is directed further radially inward again, i.e. toward the turbine longitudinal axis  9 , by the outside of the combustion chamber wall  23 . The inner cooling air flow K i  is therefore directed, first of all still within the cooling air K undivided in the main deflecting region  30 , radially outward in a path curved approximately in a C shape and is decelerated in the process and then directed radially inward in a path curved in the opposite direction approximately in a C shape. Overall, the flow through the diffuser  27  and further into the inner flow transfer space  38  approximately describes a double S-shaped path. The radii of curvature within this path are sufficiently large in order to cause only small energy losses during the flow. 
     Furthermore, baffle elements or fastening elements  41  are arranged at the downstream end  31  of the diffuser  27  in both the direction of the outer flow transfer space  37  and the direction of the inner flow transfer space  38 . 
     The outer cooling air flow K a  is directed radially outward, perpendicularly to the turbine longitudinal axis  9 , by the dividing fork  33 . In continuation, the outer cooling air flow K a  is directed past the outer combustion chamber shell  28  and into the wall lining space or wall cooling space  26 . Here, too, in a similar manner to the inner cooling air flow K i , the flow is directed with large radii of curvature, in the course of which no abrupt widening of cross section occurs. The combustion chamber shells  28 ,  29  are cooled from outside by the cooling air flows or partial flows K a , K i . 
     The fuel injector  10  is arranged approximately centrally in the combustion chamber rear wall  42 . A straight line running through the combustion chamber rear wall  42  encloses a wall angle βof about 45° with the turbine longitudinal axis  9 . The wall angle βthus corresponds approximately to the dividing angle α. The flow-dividing element  32  arranged obliquely relative to the turbine longitudinal axis  9  by the dividing angle a splits the main deflecting region  30  into a top sectional passage  43  and a bottom sectional passage  44 , which both have approximately the same cross section. The cooling air flow in the diffuser  27  can be divided in a specifically asymmetrical manner by a laterally offset arrangement of the flow-dividing element  32 , i.e. by an arrangement offset along the inner combustion chamber shell  29 , if, for example, the outer combustion chamber shell and the inner combustion chamber shell  29  have a different cooling air requirement. 
     The bleeding for turbine cooling air is effected by a tube  45  which projects into the bottom sectional passage  44 . The end  46  of said tube  45  is angled like a periscope, and its tube opening faces the inner air flow K i , so that some of the air flow K i  can flow as turbine cooling air into the tube  45 . At the other end of the tube  45 , the turbine cooling air flows into an annular passage  47  which extends along the rotor and directs the turbine cooling air to the turbine  6 . It is used there for cooling the moving and the guide blades  12 ,  14 .