Abstract:
A gas turbine system comprises a compressor that takes in suction air on the inlet side and compresses it to compressor end air that is available on the outlet side, a combustor in which a fuel is burned by using the compressor end air while resulting in the formation of hot gas, as well as a turbine in which the hot gas is expanded while providing work output. In a method for cooling this gas turbine system, compressed air is removed from the compressor, is fed as cooling air for cooling inside an internal cooling channel through thermally loaded components of the combustor and/or the turbine, is then recooled and subsequently compressed and added to the compressor end air. The influence of the cooling on the efficiency of the system is minimized by the fact that at least part of the compressor end air is used to recool the cooling air.

Description:
FIELD OF THE INVENTION 
     This invention relates to the field of gas turbines, and more particularly to methods and apparatus for cooling 
     BACKGROUND OF THE INVENTION 
     In order to cool their hot parts, in particular the combustor and the turbine through which the hot gas flows, existing gas turbines (gas turbine systems) use either cooling media taken from the compressor at a suitable pressure-and which sometimes are further cooled-and which, after they have been used to cool the hot parts, are added to the turbine stream; or these existing gas turbines use closed cooling circuits supplied from an external cooling medium source, in most cases water or steam. Such a method and such a gas turbine system are known, for example, from publication U.S. Pat. No. 5,611,197. In the latter case, frequently found in combination power plants, the cooling heat often can be used in the process that follows. Another possibility, described, for example, in EP-A2-0 899 425 of the applicant, combines, especially in the case of blade cooling, a closed steam cooling system in the main part of the blade with an open cooling system in the area of the leading blade edge. 
     The first category has the disadvantage that the cooling medium, which inherently bypasses heating in the combustor, in most cases undergoes a higher pressure loss in the cooling section than is necessary for the cooling task. In addition, mixing losses are created when the cooling medium enters the main stream. Both represent significant process losses that have an important adverse effect on the efficiency of the process overall. 
     The second category of externally supplied cooling systems and, in particular, also the third category of the combined cooling systems, does not have these disadvantages or is only affected by them to a limited degree; however, their operation becomes dependent on an external coolant supply, which is associated with an increased level of complexity as well as increased cost and safety risks. 
     The initially mentioned U.S. Pat. No. 5,611,197 discloses a gas turbine with a closed cooling system for the guide and rotating blades and the hot gas housing of the turbine, in which air with a specific pressure is removed from the compressor at an intermediate pressure level or at the outlet, this air is supplied as cooling air through the components to be cooled, and is then again fed into the compressor at a suitable, lower pressure level. Prior to being fed into the compressor, the returned cooling air hereby also can be additionally cooled inside a cooler. 
     This known type of closed cooling circuit has significant advantages in terms of simplicity of design and operation and influence on the overall efficiency when compared to the types of cooling described previously in this document. The disadvantage is, however, that in the case of a recooling of the cooling air, external cooling media (52 in the figure of U.S. Pat. No. 5,611,197) are used to cool down the returned cooling air in a heat exchanger (50). The heat removed in the heat exchanger in this way is removed in an efficiency-reducing manner from the process of the gas turbine system and at most can be utilized with additional expenditure. 
     SUMMARY OF THE INVENTION 
     It is therefore the objective of the invention to disclose a cooling method for a gas turbine as well as a gas turbine system for performing said method that avoids the disadvantages of known methods of gas turbine systems and is characterized, in particular, by a simple and substantially efficiency-neutral recooling. 
     The concept of the invention is to perform at least a substantial part of the recooling with at least one part of the compressor end air as a cooling medium. The heat removed from the cooling air in this way is easily returned into the process of the gas turbine system. The recooling of the cooling air with the compressor end air is hereby preferably performed in a heat exchanger, in particular, in a counter-current heat exchanger. 
     According to a first preferred embodiment of the method according to the invention, the cooling air is passed in a completely closed cooling circuit through the components to be cooled. This ensures that no compressed air passes by the combustor in an efficiency-reducing manner and reaches the main stream. 
     A second preferred embodiment is characterized in that a part of the cooling air is fed for film cooling through drilled film cooling openings on the components, in the manner of a targeted leakage, into the turbine stream. This makes it possible to achieve a very effective additional film cooling of the exterior surfaces of the components to be cooled with only slight losses of compressed air. 
     The thermally loaded components cooled with the cooling air preferably include the walls of the transition areas combustor/gas turbine and/or housing parts of the turbine and/or rotor parts of the turbine and/or blades of the turbine. If the blades of the turbine are cooled with cooling air, it is particularly effective if drilled film cooling openings are provided on the leading blade edges and/or the trailing blade edges. 
     If a pressure loss occurs in the cooling air during the cooling process, the cooling air must be recompressed after the cooling process. It is preferred that the compressor of the gas turbine system itself is used to recompress the cooling air after the cooling process, or an external compressor is used. 
     If the recooling in the heat exchanger with the compressor end air is insufficient, a further aftercooling of the cooling air is performed after the recooling with the compressor end air, for which preferably a cooler through which a separate cooling medium flows is used. However, it would also be conceivable and reasonable to inject water directly into the cooling air in order to aftercool the cooling air. 
     A preferred embodiment of the gas turbine system according to the invention has second cooling lines that merge into the compressor at an intermediate pressure level. It would also be conceivable, however, that instead of this, an external compressor is located in the second cooling lines, and that the second cooling lines merge into the outlet of the compressor of the gas turbine system. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     The invention will be described below in reference to the drawings, wherein: 
     FIG. 1 is a greatly simplified schematic diagram of a gas turbine system according to a first embodiment of the invention; 
     FIG. 2 is a cross-sectional view through a blade with film cooling at the leading blade edge and trailing blade edge, as may be connected to a cooling circuit according to FIG. 1; 
     FIG. 3 is a schematic diagram of a gas turbine system according to a second embodiment of the invention with recompression of the cooling air by an external compressor; 
     FIG. 4 is a schematic diagram of a gas turbine system according to a third embodiment of the invention with successive cooling of several rows of blades in the turbine; 
     FIG. 5 is a schematic diagram of a gas turbine system according to a fourth embodiment of the invention, in which the aftercooling of the cooling air is effected by injecting water; and 
     FIG. 6 is a schematic diagram of a gas turbine system according to a fifth embodiment of the invention, in which the walls of the combustor and/or the hot 
     FIG. 7 is a schematic diagram of a gas turbine system according to the first embodiment of the invention shown in FIG. 1, but without the presence of any leakage air so as to form a completely closed cooling circuit. gas housing of the turbine are cooled. 
    
    
     DESCRIPTION OF THE INVENTION 
     FIG. 1 shows a greatly simplified system schematic of a gas turbine system according to a first embodiment of the invention with a cooling circuit. The gas turbine system  10  comprises a (usually multi-stage) compressor  11 , a combustor  12 , and a (usually multi-stage) turbine  13 . Compressor  11  and turbine  13  are provided with corresponding rows of blades arranged on a common rotor. The compressor  11  takes in suction air  14  on the inlet side, compresses it, and outputs it on the outlet side in the form of compressor end air  15  to the combustor  12 , where it is used as combustion air for burning a (liquid or gaseous) fuel F. The hot gas  16  created during combustion is expanded in the downstream turbine  13  while providing work output, and is then passed on in the form of waste gas  35  to a chimney or—in a combination power plant—to downstream waste heat steam generator. 
     Inside the turbine  13  are provided—surrounded by a hot gas housing—various rows of guide and rotating blades that are exposed to the hot gas  16  coming from the combustor  12 , whereby the closer the blades and housing parts are located relative to the inlet of the turbine  13 , the greater the thermal load on said blades and housing parts. Given the high hot gas temperatures required for good efficiency, these thermally severely loaded components must be cooled in order to achieve a sufficient life span. 
     According to the invention, the thermally loaded components are then cooled with cooling air removed from the compressor  11  at a predetermined pressure level, are fed via a first cooling line  17  to the component to be cooled, are used for cooling there, and are then returned for the most part via a second cooling line  17 ′ to the compressor  11  and fed into it again at a lower pressure level. This type of return makes it possible for the compressor  11  to compensate for the pressure loss created during the cooling process. The cooling air, therefore, completely or at least for the most part, takes part in the combustion process as combustion air and therefore results only in small efficiency losses. The cooling of the thermally loaded components is exclusively or substantially an internal cooling, whereby the cooling air flows through cooling channels inside the components. This results in a completely or substantially closed cooling circuit. 
     The cooling circuit is not completely closed if an additional external cooling, in the form of a film cooling, is provided or if intentional or unintentional leakages do occur. For this purpose, for example, outflow openings (drilled film cooling openings) are provided on the component to be cooled, through which openings a part of the circulating cooling air flows to the outside in the form of leakage air  18  and forms a cooling film on the hot gas-loaded external surface of the component. The content of leakage air  18  is hereby selected so that on the one hand the overall efficiency of the system is only slightly reduced, while on the other hand an effective film cooling is achieved. The leakage air  18  that flows into the turbine stream and therefore can no longer be passed through the combustor  12  is symbolized in FIG. 1 by small arrows  18  extending from the cooling circuit outward. 
     FIG. 7 shows a schematic of a gas turbine system according to the first embodiment of the invention as described above, but with a completely closed cooling circuit as a result of the absence of any leakage air  18 . 3   
     According to the invention, the heat absorbed by the cooling air during the cooling process then can be removed again from the cooling air and returned into the process, prior to being returned into the compressor  11 , in that, for the recooling, a heat exchanger  19 , preferably a counter-stream heat exchanger, through which at least part of the compressor end air  15  flows, is provided in the second cooling line  17 ′. The portion of compressor end air  15  that is supposed to absorb heat in the heat exchanger  19  can be adjusted with a control valve  19   a . If a further aftercooling is needed, an additional cooler  20  that works with a separate cooling medium, for example, water or steam, is provided downstream from the heat exchanger  19 . 
     The aftercooling with the cooler  20  at the same time can be used in the manner of an intermediate cooler to reduce the temperature of the air compressed in the compressor  11 . If the cooling air in the cooler  20  is recooled significantly more than would correspond to the heat uptake during the cooling process, the compressor end temperature, i.e., the temperature of the compressor end air  15 , can be lowered, which enables an increase in the pressure ratio and therefore an increase in the efficiency. 
     If the component to be cooled is a blade or row of blades of the turbine  13 , the leakage air  18 —if the cooling circuit is not completely closed—is preferably used to cool the leading blade edges and/or trailing blade edges of the blade(s) by film cooling. A cross-section of an exemplary blade  23  suitable for this purpose is shown in FIG.  2 . The blade  23  has a pressure-side blade wall  24  and a suction-side blade wall  25  that both merge at the leading blade edge  21  and the trailing blade edge  22 . Inside the blade  23 —separated by support walls from each other-various cooling channels  26 , . . . ,  30  that extend in axial direction of the blade  23  (i.e., vertical to the drawing plane) are provided; the cooling air flows through these cooling channels in alternating direction (see, for example, EP-A2-0 899 425). From the cooling channels  28  and  30  located in the area of the edges  21 ,  22 , drilled film cooling openings  33  or  34 , through which the leakage air  18  is able to flow out and form a cooling film on the outside, extend towards the outside (also see, for example, US-A-5,498,133). The cooling channels  28 ,  30  are hereby supplied with cooling air from the adjoining cooling channels  27 ,  29  via connecting channels  31 ,  32 . 
     Based on the basic schematic of the cooling system according to the invention as shown in FIG. 1, different variations that are adapted to different applications and thus have specific advantages can be realized. In the exemplary embodiment of a gas turbine system  36  shown in FIG. 3, one of these variations is realized. In the cooling circuit shown here, formed by cooling lines  17  and  17 ′, the compressor end air  15  with the compressor end mass stream mv is divided into three partial streams with the mass streams m 1 , m 2 , and m 3 , whereby mv =m 1  +m 2  +m 3 , and each one of the partial mass streams is ≧0. The first partial mass stream ml reaches the combustor  12  directly. The second partial mass stream m 2  flows through the cooling lines  17  and  17 ′ and the heat exchanger  19  in order to cool the turbine  13  and is then recompressed by an external compressor  37 . The third partial mass stream m 3  and the recompressed second partial mass stream m 2  flow in counter-current through the heat exchanger  19 , and these two mass streams are combined downstream from the heat exchanger and fed together with the first partial mass stream ml to the combustor  12 . The necessary recompression after passing through the heat exchanger  19  is therefore performed not in the compressor  11  of the gas turbine system  35 , but rather in the external compressor  37 . Here also an additional cooler can be provided for aftercooling. If the cooling air in this arrangement is brought by the external compressor  37  to a pressure that is higher than the pressure of the compressor end air  15 , it is possible and advantageous to use the compressed cooling air for a showerhead cooling in a first turbine stage of the turbine  13 . It is, however, also conceivable and reasonable to use a partial mass stream, such as m 2 , to cool parts of the combustor, as is explained in more detail below for a comparable solution in reference to FIG.  6 . 
     FIG. 4 shows another embodiment of the cooling system according to the invention. The cooling circuit of the gas turbine system  38  with cooling lines  17  and  17 ′ in this example is used not only for a single row of blades of the turbine  13 , but for several rows of blades  39 ,  40 , and  41 , through which the cooling air flows sequentially. In each of the rows of blades  39 , . . . ,  41 , leakage air  18  again can flow into the main stream of the turbine  13  in order to film-cool the edges. 
     Another possibility for aftercooling is shown in the embodiment in FIG.  5 . In the gas turbine system  42  of this figure, an injection device  43  is inserted into the cooling circuit with the cooling lines  17 ,  17 ′ downstream from the heat exchanger  19  for aftercooling. Analogously to a type of “quench cooling”, water is injected here into the cooling air. The temperature reduction of the cooling air that can be achieved with this is preferably designed so that the temperature of the mixed gas is reduced after the recooled cooling air is mixed with the main air flowing through the compressor  11 . As already mentioned above, this makes it possible to increase the system&#39;s efficiency. 
     Finally, according to FIG. 6, it is possible within the scope of this invention that in a gas turbine system  44 , instead of or in addition to the blades of the turbine  13 , other components of the system with high thermal loads are cooled with air in the closed circuit. In FIG. 6, for example, the cooling circuit with cooling lines  45 ,  45 ′ and the heat exchanger  19  is designed for cooling the walls of the combustion chamber  12  or the combustion chamber liners by way of an internal cooling air circulation and external film cooling with leakage air  18 . Another cooling circuit (drawn with broken lines) with cooling lines  46 ,  46 ′ ensures an internal and, if needed, external cooling of the hot gas housing of the turbine  13 , in particular in the inlet area of the hot gasses. 
     Overall, the invention provides an effective cooling of the thermally loaded components of a gas turbine system, which is simple in its design and operation, can be used flexibly, and has only minor effects on the overall efficiency of the system.