Patent Publication Number: US-6217280-B1

Title: Turbine inter-disk cavity cooling air compressor

Description:
RELATED APPLICATION 
     This application is a continuation-in-part of commonly assigned patent application Ser. No. 08/730,720, filed Oct. 7, 1995, now U.S. Pat. No. 5,704,764, issued Jan. 6, 1998. 
    
    
     The United States Government has rights in this invention pursuant to Contract No. DE-AC21-93M30247 with the Department of Energy. 
    
    
     FIELD OF THE INVENTION 
     The invention relates to cooling systems for use with turbine engines. More particularly, the invention relates to pressure changing mechanisms disposed in turbine engines for changing the pressure of a cooling medium that is flowing through the turbine engine. This invention also relates to methods and systems that employ the pressure changing mechanisms. 
     BACKGROUND OF THE INVENTION 
     Pressurized air is among the more common cooling mediums used to cool various components in gas turbine engines. Generally in such systems, compressed air is drawn from the combustor shell and traverses a closed loop cooling system in which the air cools components of the turbine engine, e.g., the vanes, the blades and the combustors. Typically, the air is first filtered and cooled before its use as a coolant. After being cooled, the air is directed to the components to be cooled, and then the air is returned to the discharge of the compressor or the combustor shell of the gas turbine engine. 
     In such a closed loop system, the cooling air must be sufficiently pressurized in order to re-enter the combustor shell or mix with the air discharged from the compressor. Unfortunately, within the cooling circuit, the air generally experiences a pressure loss. This pressure loss is caused by the resistance of bends, orifices and other obstructions. To overcome these pressure drops and increase the pressure of the cooling medium to approximately that of the combustor shell or the discharge of the compressor, the air coolant, in some applications, is routed out of the turbine engine to an external compressor before it is returned to the combustion turbine and eventually the combustor shell. In the external compressor, the air coolant may be compressed about 60 PSI. Significantly, external compressors are expensive components, with costs in the $300,000 range. Other costs are associated with the use of external compressors, e.g., back up compressors, piping, operation, maintenance, floor space and the like. Applicants have recognized that a turbine engine that internally provides the pressurization required for the air coolant to reenter the combustor shell would eliminate the need for external compressors, thereby providing substantial economic benefits. 
     Thus, there is a need for pressure changing mechanisms that function within turbines and compress the cooling medium and thereby eliminate the need for external compressors. There is also a need for improved systems and methods for using the pressure reducing mechanisms that operate within turbines. 
     SUMMARY OF THE INVENTION 
     A cooling circuit for a combustion turbine may include a compressor, a combustor shell, an external cooler, a component of the turbine to be cooled and the pressure changing mechanism of this invention. As is conventional with cooling circuits, a cooling medium flows from the compressor to the combustor shell. From the combustor shell, the cooling medium flows through the external cooler and the component of the turbine to be cooled. After flowing through the component of the turbine to be cooled the pressure of the cooling medium is less than that of the combustor shell. Therefore, in order to return the cooling medium to the combustor shell, the pressure of the cooling medium must be raised. This invention includes a pressure changing mechanism disposed within the combustion turbine and within the cooling circuit of the combustion turbine that increases the pressure of the cooling medium from the pressure at which it exits the component to be cooled to approximately the pressure of the combustor shell. 
     According to one aspect of this invention, the combustion turbine components that are cooled by the cooling circuit are the rotating blades disposed within the turbine section of the combustion turbine. In order to cool the rotating blades, the cooling circuit further includes a flow path defined within the rotating disks of the combustion turbine. In this type of cooling circuit, the cooling medium flows from the combustor shell through the rotating disks and then through the rotating blades. From the rotating blades, the cooling medium then flows through the pressure changing mechanism. 
     In a preferred embodiment of this invention the pressure changing mechanism includes a plurality of diffusing vanes that are disposed circumferentially around a torque tube casing of the combustion turbine. A diffusing channel is defined between every two diffusing vanes. The diffusing channels receives cooling medium after it has flowed through the rotating disks and blades of the combustion turbine. Preferably, the geometry of each of these diffusing vanes and channels is such that when it receives the cooling medium from each of the rotating blades and disks it increases the pressure of the cooling medium to approximately that of the combustor shell. 
     In order to change the pressure of the cooling medium, each diffusing vane is preferably curved and each of the diffusing channels has a portion of which has an increasing cross-sectional area. When the cooling medium enters the channels defined by the diffusing vanes it slows down in this portion of the channel that has an increasing cross sectional area. Because of this deceleration and subsequent decrease in velocity, the static pressure of the cooling medium is increased to approximate that of the combustor shell pressure. 
     In another preferred embodiment of this invention, the pressure changing mechanism includes a ring disposed within the torque tube casing of the combustion turbine. Defined within the ring are a plurality of diffusing channels. These channels are disposed between the rotating blades and the combustor shell so that the cooling medium flows from the rotating blades through the diffusing channels and into the combustor shell. Each of these diffusing channels has a geometry that causes the pressure of the cooling medium to approximate that of the combustor shell. Preferably, each of these diffusing channels has a receiving end and an exhausting end. The receiving end receives coolant from the rotating disks and blades, and the exhausting end exhausts cooling medium to the combustor shell. The cross-sectional area of the receiving end is preferably smaller than the cross-sectional area of the exhausting end, so that the cooling medium diffuses within the diffusing channel, and the static pressure of the cooling medium is thereby increased to approximately that of combustor shell pressure. 
     Other features of this invention are described below. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     The foregoing summary, as well as the following detailed description of the preferred embodiments, is better understood when read in conjunction with the appended drawings. For the purpose of illustrating the invention, there is shown in the drawings an embodiment that is presently preferred, it being understood, however, that the invention is not limited to the specific methods and instrumentalities disclosed. 
     In the drawings: 
     FIG. 1 is a sectional view of a turbine section of a turbine engine wherein the present invention may be employed; 
     FIG. 1A is a sectional view of a portion of a gas turbine engine showing a portion of the air coolant path; 
     FIG. 2 is a front view of a portion of a rotor disk employing aspects of the present invention; 
     FIG. 3 is a sectional view of a presently preferred embodiment of the present invention within the rotor disk inter-cavity wherein the geometric shape of the ridges is rectangular; 
     FIG. 4 is a sectional view of a presently preferred embodiment of the present invention within the rotor disk inter-cavity wherein the geometric shape of the ridges is triangular; 
     FIG. 5 is a schematic diagram of a preferred embodiment of the system of this invention; 
     FIG. 6 is a diagrammatical view of a preferred embodiment of a portion of the system of FIG. 5; 
     FIG. 7 is a diagrammatical view of a portion of the system of FIG. 5; 
     FIG. 8 is a cross-sectional view along line  8 — 8  of FIG. 7; 
     FIG. 8A is an enlargement of the cooling air velocity vector diagram shown in FIG. 8; 
     FIG. 9 is a diagrammatical view of a preferred embodiment of a portion of the system of FIG. 5; 
     FIG. 10 is a top view of a preferred embodiment of this invention; 
     FIG. 11 is a cross-sectional view of the preferred embodiment of FIG. 10; 
     FIG. 12 is a cross-section taken along line  12 — 12  of FIG. 10; 
     FIG. 13 is a cross-section taken along line  13 — 13 ; of FIG. 10; and 
     FIG. 14 is a cross-section taken along line  14 — 14  of FIG.  10 . 
    
    
     DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS 
     Referring now to the drawings wherein like numerals indicate like elements throughout, FIG. 1 presents a diagram of a turbine  10  portion of a gas turbine or combustion engine wherein the present invention may be employed. As shown in FIG. 1, the turbine  10  comprises a plurality of turbine rotor disks  12 . These rotor disks  12  are arranged in parallel planes to form a turbine shaft, which is rotatably disposed within the turbine  10 . An inter-disk cavity  16  is formed by the space between the rotor disks  12 . The rotor disks  12  can then rotate in tandem within the turbine  10 . Rotor blades  13 , which are attached to the rotor disks  12 , are disposed within the hot gas path  15 . As the hot gas expands axially through the turbine  10 , the rotor blades  13  and rotor disk  12  assembly are caused to rotate. 
     Coolant must be provided to the rotor blades  13  as well as other turbine engine components because of the exposure to extreme heat from the hot gas expanding through the turbine  10 . In a presently preferred embodiment of the present invention, the coolant comprises air; however, persons skilled in the art will appreciate that other gases or combinations of gases, such as steam, can be substituted for the air without affecting the function or novelty of the present invention. 
     Referring now to FIG. 1A, the path of the air coolant is shown as it flows through the turbine  10  to reach the rotor blades  13 . In the presently preferred embodiment, the air coolant flows through the turbine  10  from the rear of the turbine  10  toward the front of the turbine  10 . As will be explained more fully below, along the coolant flow path, a portion of the air coolant is shunted off to provide the coolant needs for each set of rotor blades  13 . 
     The air coolant flows through each rotor disk  12  via a duct  14 . In particular, the air coolant enters the last rotor disk  12   a  via duct  14   a . The air coolant then enters the inter-disk cavity  16   a . As shown, a portion of the air coolant is shunted outwardly through the inter-disk cavity  16   a  to provide the coolant needs for the rotor blades  13 . The remaining air coolant continues travels through the turbine  10  via duct  14   b  in rotor disk  12   b . After traveling through rotor disk  12   b , another portion of the air coolant enters the next inter-disk cavity  16   b . Similarly, this portion of the air is shunted outwardly to provide the cooling needs of the next set of rotor blades  13 . Subsequently, another portion of the coolant air enters the next disk  12   c  via duct  14   c.    
     As indicated above, the air coolant must be pressurized before entering the rotor blades  13 . According to the present invention, the pressurization is provided by the rotor disks  12  and the inter-disk cavity. Essentially, the air coolant enters the inter-disk cavities  16   a ,  16   b . Therein, the air pressure must be increased to provide pressure higher than compressor discharge pressure at the exit of the cooling circuit  17 . According to an aspect of the present invention, the air pressure increase is gained via the rotation of the rotor disks  12 . 
     A series of ridges  30  are disposed within the inter-disk cavities  16  to increase the pressure of the air coolant as it flows outwardly. In the presently preferred embodiments, as explained more fully below, the ridges  30  can be attached to one side of the inter-disk cavity  16 , i.e., to only one of the faces of the rotor disk  12 , or, alternatively, the ridges  30  can be attached to both sides of the inter-disk cavity  16 , i.e., both faces of the rotor disk  12 . 
     Referring now to FIGS. 2 and 3, the face of a portion of a rotor disk  12  having the ridges  30  of the present invention is depicted. In a presently preferred embodiment, spacers  20  are attached to the face of the rotor disk  12 . The spacers  20  are configured with ridges such that as each rotor disk  12  rotates about its axis  22 , the pressure of the air coolant flowing out through the inter-disk cavity outlets  24  is increased. The spacers  20  are attached to the rotor disk  12  via thru bolts  26 . As best shown in FIG. 3, each spacer  20  comprises a series of ridges  30  that extend radially outward from the center toward the periphery of the rotor disk  12 . Those skilled in the art will recognize that the length of the ridges shown in FIG. 3, although depicted with straight lines, may be a variety of shapes, such as curved lines. The cross-section of the ridges  30  shows that the ridges  30  have a rectangular cross-section. Air passages  32  remain between the ridges  30 . As the rotor disk  12  rotates about the turbine shaft, the pressure flowing through the air passages  32  is greatly increased, i.e., on the order of 50 psi. Thus, the pressure rise within the inter-disk cavity  16  approaches that of an external compressor. 
     Referring now to FIG. 4, another presently preferred embodiment of the present invention is illustrated. As shown, in this embodiment, the ridges  30   a  and  30   b  rise off of both rotor disk faces that form the inter-disk cavity to form air passages  32   a  and  32   b . Moreover, the ridges  30   a  and  30   b  are not formed of separate spacers that are attached to the face of the rotor disk  12 , but rather are formed as part of the face of the rotor disk  12 . The rotor disk  12  can be machined to create the ridges  30   a  and  30   b  with the desired cross-section or, alternatively, cast as a single rotor disk  12  having ridges  30   a  and  30   b  with the desired cross-section. 
     FIG. 5 is a schematic diagram of a preferred embodiment of a cooling circuit  17  of this invention. As shown, the cooling circuit  17  may include a compressor  34 , a combustor shell  36 , a filter  38 , an external cooler  40 , the rotating disks  12   a ,  12   b ,  12   c  of the turbine section  41  and a pressure changing mechanism  42  of this invention, which is described in more detail below. As is conventional, the compressor  34  produces compressed air that may be used as the cooling medium. This compressed air is exhausted into the combustor shell  36 . From the combustor shell  36 , the compressed air is directed to the filter  38  and the external cooler  40 . The filter  38  removes impurities from the cooling medium, and the cooling medium is cooled by the external cooler  40 . From the external cooler  40 , the cooling medium may be directed to the rotating disks  12   a ,  12   b ,  12   c  and blades  13  of the turbine  10 , where the cooling medium removes heat from the rotating discs  12   a ,  12   b ,  12   c  and blades  13 . After the cooling medium has been heated, the cooling medium may be directed to the pressure changing mechanism  42  of this invention where the pressure of the cooling medium is increased to approximate that of the combustor shell  36 . After the pressure changing mechanism  42 , the cooling medium flows to the combustor shell  36 . In the combustor shell  36 , the cooling medium mixes with air exhausted from the compressor and is directed either back through the cooling circuit  17  or to the combustor  44 , of the turbine. 
     Because the cooling medium mixes with the compressed air exhausted by the compressor  34 , it must have a pressure that approximates that of the air in the combustor shell  36 . In a preferred embodiment of this invention, the pressure changing mechanism  42  increases the pressure of the cooling medium exhausted from the rotating disks  12   a ,  12   b ,  12   c  and blades  13  to approximate the pressure of the air in the combustor shell  36 . In a preferred embodiment of this invention, the pressure changing mechanism  42  is employed in a 501 Advanced Turbine System (ATS) manufactured by Westinghouse Electric Corporation. In this embodiment, the pressure changing mechanism  42  increases the pressure of the cooling medium about 25 psi. to a pressure of about 395 psia. Since the pressure of the combustor shell  36  is preferably about 390 psia., the pressure of the cooling medium after flowing through the pressure changing mechanism  42  approximates that of the combustor shell  36 . These specific pressures are not intended to be limiting, and provided by way of example and to explain the operation of a preferred embodiment of this invention. 
     According to a preferred embodiment of this invention, the turbine may have a torque tube casing  46 . This torque tube casing is stationary within the turbine. A portion of the cooling circuit  17  is defined within the torque tube casing. Preferably, the pressure changing mechanism  42  is disposed within the torque tube casing  46  so that the cooling medium can flow from the rotating disks  12   a ,  12   b ,  12   c  and blades  13  to the pressure changing mechanism  42 . After flowing through the torque tube casing  46 , the cooling medium preferably flows to the combustor shell  36  of the combustion turbine where it mixes with air that has been compressed by the compressor  34 . 
     Provided below is a description of two embodiments of the pressure changing mechanism  42  of this invention. Although two embodiments of the pressure changing mechanism  42  are provided, this invention is not limited to these two embodiments and may include other pressure changing mechanisms  42  disposed within the turbine  10  that change the pressure of the cooling medium to approximate that of the combustor shell  36 . 
     The pressure changing mechanism  42  of this invention may be employed with or without the pressurization caused by the rotating disks described above. That is, the pressure changing mechanism  42  may be employed with rotating disks that do not have the ridges  30  or the spacers  20  described above or with the described rotating disks that do have ridges  30  and/or spacers  20 . Except as indicated, the discussion provided below is with reference to a cooling circuit  17  that employs rotating disks that have spacers  20  and/or ridges  30 . 
     A preferred embodiment of the pressure changing mechanism  42  of this invention is shown in FIGS. 6-8. In this embodiment, the pressure changing mechanism  42  includes a plurality of diffusing vanes. The diffusing vanes preferably have a curved shape. Although only one diffusing vane is shown in FIG. 6, it will be understood that a combustion turbine  10  may have a plurality of similar diffusing vanes disposed circumferentially around the torque tube casing, as shown in FIG.  8 . Preferably, each diffusing vane is spaced about the same radial distance from the centerline of the combustion turbine. A diffusing channel  52  is defined in between every two diffusing vanes. Each of these channels  52  is disposed proximal to the rotating disk  12   d  so that the channels may receive cooling medium that has flowed through the rotating disks  12   a ,  12   b ,  12   c ,  12   d  and the rotating blades  13 . 
     Each of the diffusing vanes is shaped so that when the cooling medium flows through the channels  52  defined by the respective diffusing vanes the pressure of the cooling medium increases. This increase in pressure is due to the shape of the diffusing vanes and channels  52 . The diffusing vanes and channels are shaped such that at least a portion  48  of each of the diffusing channels  52  has an increasing cross-sectional area. Because of this increasing cross-sectional area portion  48  of the diffusing channels  52 , the cooling medium slows down as it travels through the diffusing channels. As the cooling medium slows down and decelerates, the pressure, and in particular the static pressure, of the cooling medium increases. Most preferably, the geometry of the diffusing vanes and channels  52  are such that the pressure of the cooling medium increases to approximate that of the combustor shell. After flowing through the diffusing channels, the cooling medium flows to the combustor shell. 
     A plurality of flow guiding blades  58  may be disposed between the rotating disks and the diffusing vanes. Preferably, these flow guiding blades  58  are disposed circumferentially around the turbine  10  in the spacer disc  60  and are approximately equally spaced from the radius of the centerline of the turbine  10 . A flow guiding passage  59  is defined in between each of the flow guiding blades  58 . Cooling medium exhausted from the rotating disk  12   d  is directed to these flow guiding passages  59 . As the cooling medium flows through the flow guiding passages, the flow guiding blades  58  direct the flow of the cooling medium exhausted from the rotating disks to the diffusing vanes. Preferably, they redirect the flow of the cooling medium from generally in the axial direction to generally in the radial direction, as shown in FIG.  7 . 
     The manner in which the pressure of the cooling medium is increased is better understood with reference to FIGS. 8 and 8A. As described above, the cooling medium travels through the cooling path defined within the rotating disks and rotating blades  13 . As the cooling medium flows through the rotating blades and disks, it removes heat from the rotating disks and rotating blades  13 . Furthermore, as the cooling medium flows through the rotating disks and blades, the cooling medium decreases in pressure due primarily to friction losses. After flowing through the cooling path defined within the disks and blades as described above, the cooling medium exits the disks and blades and flows through the flow guiding passages  59  defined by the flow guiding blades  58 . When the cooling medium exits the flow guiding blades  59 , the cooling medium exits with an absolute velocity and a relative velocity. The approximate direction of these velocities is labeled in FIG.  8 . The cooling medium is generally flowing in a radial direction as it exits the flow guiding passages  59 . As shown, the stationary diffusing vanes are constructed so that the cooling medium will flow through the diffusing channels due to the direction of its absolute velocity. While flowing through the diffusing channels, the cooling medium slows down and increases in pressure. Thus, by constructing the diffusing vanes such that the cooling medium will slow down as it exits the cooling path defined within the rotating disks, the diffusing vanes cause the pressure of the cooling medium to be increased. 
     In another preferred embodiment of this invention, the pressure changing mechanism  42  includes a ring  62  disposed within the torque tube casing that has a plurality of diffusing channels  64 . A preferred embodiment of the ring  62  and channels  64  are shown in FIGS. 9-14. Preferably, the combustion turbine has a plurality of similar diffusing channels  64  disposed circumferentially around the torque tube casing  46 . The ring  62  may be machined to form the diffusing channels  64 . As shown the diffusing channels  64  preferably have a geometry that will increase the pressure of the cooling medium exhausted from the rotating disks  12   a ,  12   b ,  12   c ,  12   d  to approximate the pressure of the combustor shell  36 . In a preferred embodiment, the diffusing channels  64  have a receiving end  66  and an exhausting end  68 . The receiving end  66  receives cooling medium exhausted from the rotating disks  12   a ,  12   b ,  12   c ,  12   d  and blades  13  and the exhausting end  68  exhausts the cooling medium after it has flowed through the diffusing channels  64  and directs the cooling medium to the combustor shell  36 . Preferably, the depth d r  of the receiving end  66  is smaller than the depth d e  of the exhausting end  68 . Moreover, it is preferable that the diffusing channels  64  have at least a portion  76  that has a gradually increasing depth or cross-sectional area. These depths or areas are selectively chosen to increase the pressure of the cooling medium from its initial pressure when it is received from the rotating disks  12   a ,  12   b ,  12   c ,  12   d  to approximately the pressure of the combustor shell  36 . In a preferred embodiment, the portion  76  of each channel  64  that has a gradually increasing depth or cross-sectional area is the entire diffusing channel  64 . 
     The diffusing channels  64  increase the pressure of the cooling medium received from the rotating disks  12   a    12   b ,  12   c ,  12   d  as follows. The cooling medium exits the rotating disks  12   a ,  12   b ,  12   c ,  12   d  and blades with a directional velocity due to the rotation of the rotating disks that is generally in the tangential direction of the turbine. This pressure at which the cooling medium exists the rotating blades and disks is less than the combustor shell pressure. From the rotating disks  12   a ,  12   b ,  12   c ,  12   d  the coolant enters the diffusing channels  64 . In the diffusing channels  64 , the cooling medium decelerates due to the geometry of the diffusing channels  64 , and the pressure of the cooling medium is thereby increased. This occurs because of the increasing cross-sectional area of the diffusing channel. Furthermore, the cooling medium changes its direction as it is funneled into the channels  64  and impinges against the channels  64 . 
     In a preferred embodiment of this invention, the dimensions of the diffusing channels are as follows. The depth d r  of the receiving end is about 13.2 mm., and the depth d e  of the exhausting end is about 40.7 mm. The depth d s  of the channel at the end of the straight portion is about 28.9 mm. Preferably, the angle A at which the centerline of the channel is disposed is about 60 degrees relative to a line L that is parallel to the centerline of the turbine. The width w of the channel is preferably about 13.2 mm., and the length D l  of the straight portion is preferably about 150.0 mm. 
     As mentioned above, the pressure changing mechanism  42  of this invention may be employed in a cooling circuit  17  that has spacers  20  and/or ridges  30 . When employed with rotating disks that have ridges  30  or spacers  20  for increasing the pressure of the cooling medium, the pressure changing mechanism  42  further increases the pressure of the cooling medium over the increase provided by the rotating disks. It will be appreciated that the geometry of either or both of the pressure changing mechanism  42  or the ridges/spacers  30 ,  20  of the rotating disks may have to be adjusted to obtain the appropriate change in the pressure of the cooling medium depending on the application in which they are employed. 
     In summary, the pressure changing mechanism  42  of this invention changes the pressure of a cooling medium flowing within a cooling circuit  17  of a combustion turbine  10 . Preferably, the pressure changing mechanism  42 , increases the pressure of the cooling medium from the pressure at which the cooling medium is exhausted from the rotating disks and blades to a pressure of the combustor shell  36  of the combustion turbine  10 . 
     By providing a pressure changing mechanism  42  within a combustion turbine  10  in combination with the spacers  20  and/or ridges  30 , the need for an external compressor in a combustion turbine cooling circuit  17  is eliminated. Potentially, this may result in economic benefits due to decreased capital expenditures and maintenance costs associated with external compressors. 
     Although the preferred embodiments of this invention are described with reference to air as the cooling medium, it will be understood that other cooling mediums may be employed with this invention and the cooling circuit  17  may vary depending on the cooling medium employed. This invention is intended to include pressure changing mechanisms  42  disposed initially within a combustion turbine  10  that change the pressure of a cooling medium as it flows through a cooling circuit  17  of a combustion turbine  10 . 
     The present invention may be embodied in other specific forms without departing from the spirit or essential attributes thereof. For example, square ridges or some other shaped ridge cross-section could be used to generate the inter-disk cavity pressure. Accordingly, reference should be made to the appended claims, rather than to the foregoing specification, as indicating the scope of the invention.