Abstract:
In gas turbines, compressed air is supplied via an air duct to combustion chambers and is heated there. Pressure losses in the air duct should be minimized in order to ensure good overall efficiency. This is achieved by the compressed air flowing with approximately constant velocity in the air duct from the compressor to the inlet into the combustion chamber. This is supported by the effective cross section of the air duct being almost constant over this distance.

Description:
[0001]    The present application hereby claims priority under 35 U.S.C. Section 119 on European Patent application number 01114599.2 filed Jun. 18, 2001, the entire contents of which are hereby incorporated by reference.  
         FIELD OF THE INVENTION  
         [0002]    The invention generally relates to a gas turbine with a compressor for air. More particularly, it relates to one which is heated in a plurality of combustion chambers connected in parallel with respect to flow, before it flows via a transfer duct to a gas duct in a turbine. It additionally can relate to a method of operating a gas turbine.  
         BACKGROUND OF THE INVENTION  
         [0003]    In gas turbines, induced air is usually compressed initially, and is then heated in combustion chambers in order to achieve an economic power density. The hot gas generated in this process then drives a turbine.  
           [0004]    In order to achieve good overall efficiency, it is inter alia necessary to keep flow losses small during the guidance of the compressed air. At the same time, however, various components of the turbine installation have to be cooled with the compressed and as yet unheated air. Thus, for example, a transfer or connecting duct, through which hot gas from the combustion chambers flows to the turbine, must be protected from overheating in order to avoid damage.  
           [0005]    An arrangement which has widespread application for this purpose is given in FIG. 1 in U.S. Pat. No. 4,719,748. In this arrangement, a long connecting duct between a combustion chamber and a turbine inlet is located directly in an air duct through which compressed air flows to a burner. In this arrangement, no diffuser is shown for air deflection and the flow velocity of the air has fallen greatly on reaching the connecting duct. In consequence, correct cooling is at best possible at relatively low temperatures of the hot gas because higher temperatures require a specific flow velocity both for the compressed air and for the hot gas and a specific air duct height and alignment. As far as can be seen, adequate cooling cannot be achieved with this solution for either the upper side or the lower side of the connecting duct because, on the one hand, the volume of the air duct is very large in this region and because, in addition, both the length of the duct section to be cooled and the distance to be traversed by the compressed air after emergence from a compressor are relatively long.  
           [0006]    In addition, however, a complicated cooling device, in which one combustion chamber and a connecting duct leading from this to a turbine are covered by a second wall relative to the flow of the compressed air, is the subject matter of the cited U.S. Pat. No. 4,719,748 in FIGS.  2  to  7  and the associated description. A multiplicity of openings, through which the compressed air is specifically deflected onto the wall sections to be cooled, are provided in this second wall. Although good cooling can be achieved by the variations given for this solution with respect to the number, the size and the shape of these openings, a disadvantage of this arrangement is a not insubstantial, unavoidable pressure loss in the compressed air because the latter must be repeatedly decelerated and accelerated again.  
         SUMMARY OF THE INVENTION  
         [0007]    An embodiment of the invention includes an object of creating an arrangement, for a gas turbine, in which an unavoidable pressure loss in the flow of the compressed air is further reduced.  
           [0008]    This object may be achieved, for example, by the compressed air flowing with approximately constant velocity over the whole distance in an air duct from the outlet of the compressor to the inlet into the combustion chambers. In this arrangement, the transfer duct may be expediently shorter than the diameter dimension of one of the combustion chambers. This solution is surprisingly advantageous because not only the pressure drop in the air duct but, in addition, a pressure drop in the transfer duct also are lowered to a very small value. In this arrangement, a constant velocity of the air in the air duct may be achieved by the effective cross section of the air duct being almost constant over the whole distance from the outlet of the compressor to the inlet into the combustion chambers. 
       
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0009]    An exemplary embodiment of the invention is explained in more detail using drawings, wherein:  
         [0010]    [0010]FIG. 1 shows an excerpt from a gas turbine in longitudinal section,  
         [0011]    [0011]FIG. 2 shows a section along the line II-II in FIG. 1,  
         [0012]    [0012]FIG. 3 shows a section along the line III-III in FIG. 1, and  
         [0013]    [0013]FIG. 4 shows a view in the direction IV of FIG. 2 onto an outer casing (not shown there) of a combustion chamber. 
     
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS  
       [0014]    A rotor  1 , shown as an excerpt, of a gas turbine installation rotates about a center line  2 . In a compressor  3 , compressed air leaves the compressor  3  through a ring of guide vanes  4  and flows, in the direction of the arrows  5 , initially through a duct section  6 , which is parallel to the center line and circular in cross section, of an air duct which is bounded on the inside by a wall  38  and on the outside by a wall  39 .  
         [0015]    At the end of this duct section  6 , the compressed air passes struts  7 . The struts  7  support a C-shaped cross section annular deflector  8  and are anchored in the end of the duct section  6  via struts  7 . An arm  9 , which is located in the end of the duct section  6 , of the cross section of the deflector  8  forms, via its edge  9  facing upstream, a wavy line  37  oscillating about a circle concentric with the center line  1 . The wall thickness of the deflector  8  increases strongly, starting from the edge  9  and extending to its center, and is not constant in the peripheral direction of the deflector  8  either, but increases and decreases in wave form.  
         [0016]    Combustion chambers  10  for heating the compressed air are arranged radially above the deflector  8 . A cross-sectional arm, which is located radially on the outside, of the deflector  8  is essentially matched to the contour of the combustion chambers and forms, with its free end, a wave-shaped edge  35 . This outer cross-sectional arm of the deflector  8  is, in addition, also wave-shaped per se, the waves formed in this way being opposite to the waves of the wavy line  37 , as can be seen particularly well from FIG. 3.  
         [0017]    The particular shape of the deflector  8 , with its C-shaped cross section arms forming waves  35  and  37  in its peripheral direction, forces an airflow distribution in its region into a partial flow to the upper surface of the combustion chambers  10  and into a partial flow Sb to the lower surface of the combustion chambers  10 . In this arrangement, the upper surface of the combustion chambers  10  is located, relative to the gas turbine, radially on the outside and, correspondingly, the lower surface is located radially on the inside. The path distances of the partial flows and are approximately equally large, so that all parts of the cooling air have to traverse equally long paths from the compressor  3  to the inlet into the combustion chambers  10 .  
         [0018]    Each of the combustion chambers  10  is supported, from the inside, via struts  11  on an outer casing  12 , which is the outer wall of an air duct  20  and simultaneously represents a continuation of the air duct  6  for the air flowing in the direction of the arrows  5 . The casing  12  supports, on its outer free end, a cap  13  through which the air is guided into the internal space of the combustion chamber  10 .  
         [0019]    In the peripheral direction, the combustion chambers  10  are so tightly arranged adjacent to one another that the outer casings  12  have to mutually penetrate at their end facing toward the rotor  1 . In order, nevertheless, to be able to push the combustion chambers  10 , including their outer casings  12 , as far as is desired in the direction toward the rotor  1 , recesses  40  (FIG. 4) are provided on the outer casings  12 , in the region of which recesses adjacent combustion chambers  10  have a common air duct  20  between them.  
         [0020]    Fuel, for example a combustible gas or atomized, liquid fuel is, furthermore, supplied through a nozzle (not shown) to the internal space of the combustion chambers  10 , the air in the combustion chamber  10  being heated to form a hot gas  34  by the combustion of this fuel.  
         [0021]    The combustion chamber  10  and the outer casing  12  holding it are carried in a connecting piece  14  in a housing shell  15  and are fixed onto the outer end of the connecting piece  14  via a flange  16  firmly connected to the outer casing  12 . An inner end  36  of the combustion chamber  10  is located, in a sealed manner, in a transfer duct  17 , which connects the outlet of the combustion chamber  10  to a circular cross section gas duct  18  in a turbine. In order to admit hot gas  34  as evenly as possible to the gas duct  18  over its periphery, a multiplicity of, for example, ten to thirty combustion chambers  10  are evenly distributed over the periphery of the turbine installation and their openings into the transfer duct  17  are connected to one another by a peripheral duct  19  open in the direction of the gas duct  18 . The transfer duct  17  is anchored to a guidance part  22  of the turbine by thin struts  21 .  
         [0022]    In order to transfer the compressed air flowing in the direction of the arrows  5  with as little loss as possible from the duct section  6  into the ducts  20  enveloping the combustion chambers  10 , the deflector  8  supports a cross-sectional arm pointing in the direction of the free end of the combustion chambers  10 . Its edge  35  follows, in wave shape and at a small distance, the contour of the transfer duct  17  and the contours of the ends  36  of the combustion chambers  10  opening into the latter. In this way, the airflow from the duct section  6  is deflected by more than 90° into a direction parallel to the center lines of the combustion chambers  10 . By this, the combustion chambers  10  can be positioned with their center lines strongly inclined relative to the center line  1  without particular disadvantages, in which arrangement their compressor ends include an acute angle, so that they are located on a conical envelope concentric with the center line  2 .  
         [0023]    The guidance part  22  and a guidance part  23  are carried in a housing shell  24  and are secured against rotation by locking blocks  25 . On the other hand, however, the guidance parts  22  and  23  can be displaced—by, for example, hydraulic or pneumatic motors  26 —parallel to the center line over small distances, a flange  27  being elastically deformed and the deformation energy stored in it being used for restoring the guidance parts  22  and  23 . A volume enclosed by the housing shells  15  and  24  is subdivided into chambers by partitions  28 .  
         [0024]    The guidance parts  22  and  23  have a funnel-type design and support guide vanes  30 , which are fastened on their inside in guide rings  29 , the ends of the guide vanes  30  opposite to the guide rings  29  being firmly connected together by rings  31 . A ring of rotor blades  32 , which are splined onto the rotor  1  and whose free tips are opposite to guide rings  33 , is respectively provided between mutually adjacent rings of guide vanes  30 . In this arrangement, the guide rings  29  and  33  form an outer boundary to the gas duct  18  in the turbine for the hot gas  34  and the rings  31 , together with the roots of the rotor blades  32 , form an inner boundary.  
         [0025]    Parts of the turbine installation immediately exposed to the hot gas  34  are usually cooled, via ducts (not shown), by air tapped from the compressor or from the duct section  6 . In particular applications, pockets immediately bordering the transfer duct  17  and located in a dead angle of the airflow near the deflector  8  are, where necessary, also cooled in this way. These pockets are then expediently separated from the air duct by partitions (not shown) so that their free and effective cross section can be more precisely matched, in the region of the transfer duct  17 , to the cross section of the duct section  6  or the sum of the individual cross sections of the ducts  20 . This cross section can, in addition, be adjusted precisely by variation of the wall thickness of the deflector  8  both in its peripheral direction and in its cross section.  
         [0026]    Because the cross section of the duct section  6  and the sum of the individual cross sections of the ducts  20  are at least approximately equally large, a constant, equally large flow velocity is ensured for the compressed air in these duct sections. This flow velocity is maintained by the special shape of the C-shaped cross section deflector  8  even during the deflection of the compressed air by more than 90°. This avoids decelerations and renewed accelerations of the compressed air and, in consequence, losses caused by this are greatly reduced.  
         [0027]    The invention being thus described, it will be obvious that the same may be varied in many ways. Such variations are not to be regarded as a departure from the spirit and scope of the invention, and all such modifications as would be obvious to one skilled in the art are intended to be included within the scope of the following claims.