Abstract:
A disc seal assembly for use in a turbine engine. The disc seal assembly includes a plurality of outwardly extending sealing flange members that define a plurality of fluid pockets. The sealing flange members define a labyrinth flow path therebetween to limit leakage between a hot gas path and a disc cavity in the turbine engine.

Description:
This invention was made with U.S. Government support under Contract Number DE-FC26-05NT42644 awarded by the U.S. Department of Energy. The U.S. Government has certain rights to this invention. 
    
    
     FIELD OF THE INVENTION 
     The present invention relates generally to a disc seal assembly for use in a turbine engine, and more particularly, to a disc seal assembly including a plurality of sealing flange members that define a labyrinth flow path to limit leakage between a disc cavity and a hot gas passage in the turbine engine. 
     BACKGROUND OF THE INVENTION 
     In multistage rotary machines used for energy conversion for example, a fluid is used to produce rotational motion. In a gas turbine engine, for example, a gas is compressed in a compressor and mixed with a fuel source in a combustor. The combination of gas and fuel is then ignited for generating combustion gases (hot gas) that are directed to turbine stage(s) to produce rotational motion. Both the turbine stage(s) and the compressor have stationary or non-rotary components, such as vanes, for example, that cooperate with rotatable components, such as rotor blades, for example, for compressing and expanding the operational gases. Many components within the machines must be cooled by cooling air to prevent the components from overheating. 
     Cooling air and hot gas leakage between a hot gas path and a disc cavity in the machines reduces performance and efficiency. Cooling air leakage from the disc cavities into the hot gas path in airfoil channels can disrupt the flow of the hot gas and increase heat losses. Further, as more cooling air is leaked into the hot gas path, the higher the primary zone temperature in the combustor must be to achieve the required engine firing temperature. Additionally, hot gas leakage into the disc cavities yields higher disc and blade root temperatures and may result in reduced performance and reduced service life and/or failure of the components in the disc cavities. 
     In view of higher pressure ratios and higher engine firing temperatures implemented in modern machines, it is increasingly important to limit leakage between the hot gas path and the disc cavity in the machines to maximize performance and efficiency thereof. 
     In view of the foregoing considerations it would be desirable to provide a seal arrangement for use in a rotary machine, whereby the placement and configuration of sealing flanges in the arrangement limits leakage between the hot gas path and the disc cavity to thereby improve performance and efficiency of the rotary machine. 
     SUMMARY OF THE INVENTION 
     In accordance with a first aspect of the present invention, a seal assembly is provided for limiting gas leakage between a hot gas path and a disc cavity in a turbine engine comprising a plurality of stages, each stage comprising a plurality of stationary components connected by an annular inner shroud and a rotating disc supporting a plurality of blades. The seal assembly comprises a wing member extending axially from a side of the disc toward a radial surface of the annular inner shroud. The wing member includes an inner side and an outer side and a first wing flange extending radially outwardly from the outer side of the wing member toward an axial surface of the inner shroud. A first shroud flange extends radially inwardly from the axial surface of the inner shroud toward the outer side of the wing member to form, with the first wing flange, a labyrinth path between the hot gas path and the disc cavity. 
     In accordance with a second aspect of the present invention, a seal assembly is provided for limiting gas leakage between a hot gas path and a disc cavity in a turbine engine comprising a plurality of stages, each stage comprising a plurality of stationary components connected by an annular inner shroud and a rotating disc supporting a plurality of blades. The seal assembly comprises a wing member extending axially from a side of the disc toward a radial surface of the annular inner shroud. The wing member includes an inner side and an outer side, and a first wing flange extending radially outwardly from the outer side of the wing member toward an axial surface of the inner shroud. A second wing flange extends radially inwardly from the inner side of the wing member opposite from the first wing flange. A first shroud flange extends radially inwardly from the axial surface of the inner shroud toward the outer side of the wing member to form, with the first wing flange, a labyrinth path between the hot gas path and the disc cavity. 
     In accordance with a third aspect of the present invention a seal assembly is provided for limiting gas leakage between a hot gas path and a disc cavity in a turbine engine comprising a plurality of stages, each stage comprising a plurality of stationary components connected by an annular inner shroud and a rotating disc supporting a plurality of blades. The seal assembly comprises a wing member extending axially from a side of the disc toward a radial surface of the annular inner shroud. The wing member includes an inner side and an outer side, and a first wing flange extending radially outwardly from the outer side of the wing member toward an axial surface of the inner shroud. The first wing flange is curved extending in the radial direction and having a concave side facing the disc. A second wing flange extends radially inwardly from the inner side of the wing member opposite from the first wing flange. The second wing flange is curved extending in the radial direction and having a concave side facing the disc. A first shroud flange extends radially inwardly from the axial surface of the inner shroud toward the outer side of the wing member to form, with the first wing flange, a labyrinth path between the hot gas path and the disc cavity, wherein the first shroud flange includes a lip member extending axially from a distal end of the first shroud flange toward the first wing flange. A second shroud flange extends axially from the radial surface of the inner shroud toward the disc at a radial location generally between the first wing flange and the second wing flange. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       While the specification concludes with claims particularly pointing out and distinctly claiming the present invention, it is believed that the present invention will be better understood from the following description in conjunction with the accompanying Drawing Figures, in which like reference numerals identify like elements, and wherein: 
         FIG. 1  is a diagrammatic sectional view of a portion of a gas turbine engine including a disc seal assembly in accordance with the invention; 
         FIG. 2  is an enlarged sectional view of the disc seal assembly illustrated in  FIG. 1 ; and 
         FIG. 3  is an enlarged sectional view of a disc seal assembly in accordance with another embodiment of the present invention. 
     
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     In the following detailed description of the preferred embodiments, reference is made to the accompanying drawings that form a part hereof, and in which is shown by way of illustration, and not by way of limitation, specific preferred embodiments in which the invention may be practiced. It is to be understood that other embodiments may be utilized and that changes may be made without departing from the spirit and scope of the present invention. 
     Referring to  FIG. 1 , a portion of a turbine engine  10  is illustrated diagrammatically including adjoining stages  12 ,  14 , each stage comprising an array of stationary components, illustrated herein as vanes  16 , supported on inner shrouds  17 , and an array of rotating blades  18  supported on platforms  40  mounted to rotor discs  20 . The vanes  16  and the blades  18  are positioned circumferentially within the engine  10  with alternating vanes  16  and blades  18  located in the axial direction of the engine  10 . The rotor discs  20  are secured to adjacent discs  20  with spindle bolts  22 . The vanes  16  and the blades  18  extend into an annular gas passage  24 , and hot gases directed through the gas passage  24  flow past the vanes  16  and the blades  18  to remaining rotating elements. 
     First disc cavities  26  and second disc cavities  28  are illustrated located radially inwardly from the gas passage  24 . Purge air is provided from cooling gas passing through internal passages (not shown) in the vanes  16  and inner shrouds  17  to the disc cavities  26 ,  28  to cool the blades  18 . The purge air also provides a pressure balance against the pressure of the hot gases flowing in the gas passage  24  to counteract a flow of the hot gases into the disc cavities  26 ,  28 . In addition, interstage seals comprising labyrinth seals  30  may be supported at the radially inner side of the inner shrouds  17  and are engaged with surfaces defined on paired annular disc arms  32 ,  34  extending axially from opposed portions of adjoining discs  20 . An annular cooling cavity  36  is formed between the opposed portions of adjoining discs  20  on an inner side of the paired annular disc arms  32 ,  34 . The annular cooling cavity  36  receives cooling air passing through disc passages (not shown) to cool the discs  20 . 
     Structure on the discs  20  and the inner shrouds  17  cooperate to form annular disc sealing assemblies  38  between the gas passage  24  and the disc cavities  26 ,  28 , as more clearly shown in  FIG. 2 . For exemplary purposes, only one disc sealing assembly  38  formed between the gas passage  24  and the first disc cavity  26  will be described. However, it is understood that the other disc sealing assemblies  38  formed between the gas passage  24  and other disc cavities  26 ,  28  within the engine  10  are generally identical to or are substantially mirror images of the disc sealing assembly  38  described. 
       FIG. 2  shows an enlarged view illustrating the disc sealing assembly  38 . A wing member  44  extends axially from a first side  46  of the disc  20  toward a radial surface  48  of the inner shroud  17 . In the embodiment shown, the wing member  44  is formed from a high temperature alloy, such as for example an INCONEL® alloy (INCONEL® is a registered trademark of Special Metals Corporation), although any suitable material may be used to form the wing member  44  as desired. 
     Although only a single wing member  44  is shown, it should be understood that a plurality of wing members  44  may be employed to form the disc sealing assembly  38  as desired. If multiple wing members  44  are used to form the disc sealing assembly  38 , the wing members  44  are preferably located adjacent to each other extending circumferentially about the disc  20 , and the wing members  44  may include cooperating ramped or angled overlapping edges (not shown) to reduce spacing between adjacent wing members  44  and provide a sealing interface for restricting passage of gases between adjacent wing members  44 . 
     The wing member  44  includes an outer side  50  facing radially outwardly from the wing member  44  and an inner side  52  facing radially inwardly from the wing member  44 . The outer side  50  and inner side  52  may be generally arcuate shaped in the circumferential direction to substantially correspond to the arcuate shape of the disc  20  when viewed axially. 
     A first wing flange  54  extends radially outwardly from the outer side  50  of the wing member  44  toward an axial surface  56  of the inner shroud  17 . The axial surface  56  of the inner shroud  17  is located adjacent to and extends in a transverse direction from the radial surface  48  of the inner shroud  17 . In the embodiment shown, the first wing flange  54  is formed from a high temperature alloy, such as an INCONEL® alloy, for example, although any suitable material may be used to form the first wing flange  54  as desired. The first wing flange  54  may be arcuate shaped in the circumferential direction to substantially correspond to the arcuate shape of the disc  20  when viewed axially. In addition, the first wing flange  54  may be curved in the radial direction and include a concave side  58  facing the disc  20 . A distal end  60  of the first wing flange  54  is located adjacent to the axial surface  56  of the inner shroud  17 . 
     The inner shroud  17  includes a first shroud flange  66  that extends radially inwardly from the axial surface  56  of the inner shroud  17  toward a location adjacent the outer side  50  of the wing member  44 . The first shroud flange  66  may be arcuate shaped in the circumferential direction to substantially correspond to the arcuate shape of the inner shroud  17  when viewed axially. In the embodiment shown, the first shroud flange  66  is located at an axial location between the first wing flange  54  and the disc  20 . The first shroud flange  66  includes a lip member  68  that extends axially from a distal end  70  of the first shroud flange  66  toward the first wing flange  54 . A first fluid pocket P 1  is formed between the first shroud flange  66  and the disc  20 . A second fluid pocket P 2  is formed between the first wing flange  54  and the first shroud flange  66 . 
     A second wing flange  62  extends radially inwardly from the inner side  52  of the wing member  44 . In the embodiment shown, the second wing flange  62  is formed from a high temperature alloy, such as an INCONEL® alloy, for example, although any suitable material may be used to form the second wing flange  62  as desired. The second wing flange  62  may be arcuate shaped in the circumferential direction to substantially correspond to the arcuate shape of the disc  20  when viewed axially. In addition, the second wing flange  62  may be curved in the radial direction and include a concave side  64  facing the disc  20 . 
     It should be noted that the surfaces of the wing member  44 , including the surfaces of the first and second wing flanges  54 ,  62 , may be hardened or coated with a hard material in order to prevent or reduce abrasion and wear of these surfaces in the event that rubbing contact occurs with adjacent stationary surfaces. 
     During operation of the engine  10 , the cooling air in the disc cavity  26  is pumped radially outwardly by the rotation of the disc  20 . The curved configuration of the second wing flange  62  acts as an aerodynamic break and deflects the outward flowing disc boundary layer flow of air away from the disc  20  and forcing it to turn 180 degrees to pass around the edge of the second wing flange  62 . That is, the air of the boundary layer flow must flow in a direction radially inwardly toward the rotational axis of the disc  20  and then turn 180 degrees around the edge of the second wing flange  62  in order to flow radially outwardly past the wing member  44  along an outer convex side  65  of the second wing flange  62 . A limited gap or passage area is defined between the distal end  75  of the second shroud flange  74  and the wing flange midpoint  69  which operates to further restrict radial outward flow of cooling air from the disc cavity  26  into the third fluid pocket P 3 . 
     Once cooling air or gas passes into the third fluid pocket P 3 , it must follow a tortuous path defined by the labyrinth path  72  in order to escape into the gas passage  24 . Specifically, gas located within the third fluid pocket P 3  must pass around the distal end  60  of the first wing flange  54  and turn 180 degrees to enter the second fluid pocket P 2 , moving in a direction counter to the centrifugal outward pumping forces associated with the fluid boundary layer of the first wing flange  54 . Gas in the second fluid pocket P 2  must again turn 180 degrees to pass out of the second fluid pocket P 2  and into the first fluid pocket P 1  and the gas passage  24 . It should be noted that the lip  68  forces gas in the second fluid pocket P 2  to move toward an outwardly moving boundary layer associated with the concave side  58  of the first wing flange  54  to further counteract movement of gas from the second fluid pocket P 2  toward the gas passage  24 . It should also be understood that the restricted passages defined adjacent the distal end  60  of the first wing flange  54  and adjacent the distal end  70  of the first shroud flange  66  further act to restrict passage of gas through the labyrinth path  72  to the gas passage  24 . 
     In addition to restricting a flow of cooling air into the gas passage  24 , the sealing assembly  38  also provides a tortuous labyrinth path  72  that hot gases from the gas passage  24  must overcome in order to enter the disc cavity  26 . In addition, a pressure rise associated with the restricted seal clearances defined at the distal ends  60 ,  70 ,  75  of the first wing flange  54  and the first and second shroud flanges  66 ,  74 , respectively, further counteracts movement of the hot gases into the disc cavity  26 . 
       FIG. 3  shows an enlarged view illustrating a disc sealing assembly  138  in accordance with another embodiment of the invention, wherein corresponding structure to that described above with reference to  FIGS. 1 and 2  is identified by the same reference increased by 100. With the exception of a cover plate  147 , a first flexible seal  159 , a second flexible seal  163  and the particular structure of a portion of the a blade platform  140  associated with each of the blades  118 , the disc sealing assembly  138  is substantially identical to the disc sealing assembly  38  discussed above with reference to  FIGS. 1 and 2 . Accordingly, only these components and their associated functions will now be described. 
     The blade platform  140  supports a blade  118  thereon and includes a circumferentially extending annular groove  141  located adjacent an outer lip  143  thereof. The cover plate  147  may be provided as a cover for the axial end of the blade root of one or more blades  118  and is shown as including a radial outer edge  149 . The radial outer edge  149  is received in the annular groove  141  of the blade platform  140  and the cover plate  147  may be further mechanically secured in place, such as by clamping, peening, screwing, or other mechanical securing means, for example. It should be understood that a plurality of cover plates  147  may be provided around the circumference of the disc  120 , and that each cover plate  147  may include one or more wing members  144  to form the disc sealing assemblies  138 . The wing member  144  extends from the cover plate  147  toward a radial surface  148  of an inner shroud  117 . 
     The inner shroud  17  also includes a second shroud flange  74  that extends axially from the radial surface  48  of the inner shroud  17  toward the wing member  44 . The second shroud flange  74  may be arcuate shaped in the circumferential direction to substantially correspond to the arcuate shape of the inner shroud  17  when viewed axially. In the embodiment shown, the second shroud flange  74  is generally radially aligned with the wing member  44 , i.e., the second shroud flange  74  and the wing member  44  are located at generally the same radially location such that they are generally equidistant from a central axis of the engine  10 . Further, the second shroud flange  74  in the embodiment shown is located at a radial location generally between the first wing flange  54  and the second wing flange  62  and includes a distal end  75  located adjacent to a wing flange midpoint  69  between the first and second wing flanges  54 ,  62 . A third fluid pocket P 3  is formed by the first wing flange  54 , the inner shroud  17 , and the second shroud flange  74 . The first wing flange  54 , the first shroud flange  66 , and the lip member  68  cooperate to form a labyrinth path in the second fluid pocket P 2 , extending between the first fluid pocket P 1  and the third fluid pocket P 3  and indicated by the dashed line  72  in  FIG. 2 . 
     The first flexible seal  159  is disposed on a concave side  158  of a first wing flange  154  near a distal end  160  thereof and may be attached to the first wing flange  154 , such as by welding. In the embodiment shown, the first flexible seal  159  is formed from a high temperature alloy, such as an INCONEL® alloy, for example, although any suitable material may be used to form the first flexible seal  159  as desired. A thickness of the first flexible seal  159  in the embodiment shown is approximately 0.040 inches (approximately ⅓ of a thickness of the first wing flange  154 ), although the first flexible seal  159  may have other thicknesses as desired. The first flexible seal  159  may be arcuate shaped to substantially correspond to the arcuate shape of the disc  120  when viewed axially. In the embodiment shown, the first flexible seal  159  is curved in the axial direction and has a concave side  161  facing an axial surface  156  of the inner shroud  117 . Also in the embodiment shown, the first flexible seal  159  extends around a distal end  170  of a first shroud flange  166 , including a lip member  168 . 
     The second flexible seal  163  is disposed on a convex side  165  of a second wing flange  162 , is curved in the radial direction and extends axially toward a radial surface  148  of the inner shroud  117 . In the embodiment shown, the second flexible seal  163  is formed from a high temperature alloy, such as an INCONEL® alloy, for example, although any suitable material may be used to form the second flexible seal  163  as desired. A thickness of the second flexible seal  163  in the embodiment shown is approximately 0.040 inches (approximately ⅓ of a thickness of the second wing flange  162 ), although the second flexible seal  163  may have other thicknesses as desired. The second flexible seal  163  may be arcuate shaped to substantially correspond to the arcuate shape of the disc  120  when viewed axially. In the embodiment shown, the second flexible seal  163  has a convex side  167  facing the axial surface  156  of the inner shroud  117 . Also in the embodiment shown, the second flexible seal  163  extends into axially overlapping relationship to an inner surface  173  of a second shroud flange  174 . The reduced thickness of the first and second flexible seals  159 ,  163  relative to the respective first and second wing flanges  164 ,  162  contributes to flexing movement of the seals  159 ,  163  in response to a centrifugal force applied during rotation of the disc  120 , as is additionally described below. 
     The first wing flange  154 , the first flexible seal  159 , the first shroud flange  166 , and the lip member  168  of the first shroud flange  166  cooperate to form a first labyrinth path between a gas passage  124  and a disc cavity  126 , as indicated by the dashed line  172  in  FIG. 3 . The first wing flange  154 , the second wing flange  162 , the second flexible seal  163 , and the second shroud flange  174  cooperate to form a second labyrinth path between the gas passage  124  and the disc cavity  126 , as indicated by the dotted line  176  in  FIG. 3 . 
     It should be noted that the surfaces of the wing member  144 , including the surfaces of the first and second wing flanges  154 ,  162  and the surfaces of the first and second flexible seals  159 ,  163 , may be hardened or coated with a hard material in order to prevent or reduce abrasion and wear of these surfaces in the event that rubbing contact occurs with adjacent stationary surfaces. 
     The sealing assembly  138  operates in a manner substantially similar to that described for the sealing assembly of the first embodiment. However, the flexible seals  159 ,  163  operate to further restrict passage of gas, such as cooling air from the disc cavity  126  to the gas passage  124 . In particular, rotation of the disc  120 , and the resulting centrifugal force applied to the flexible seals  159 ,  163 , causes the flexible seals  159 ,  163  to move outwardly to locations closely adjacent to the distal end  170  of the first shroud flange  166  and the inner surface  173  of the second shroud flange  174 , respectively. Hence, the flexible seals  159 ,  163  additionally restrict the flow area for the respective labyrinth paths  172 ,  176 . It should also be noted that the flexible seal  163  provides an additional location for causing gas to change direction, i.e., 180 degrees, in order to pass between the disc cavity  126  and the third fluid chamber P 3   
     While  FIGS. 1 and 2  illustrate the wing member  44  incorporated into the sides of the discs  20  and  FIG. 3  illustrates the wing member  144  extending from the cover plate  147 , it should be understood that other configurations for supporting wing members may be provided. For example, wing members may be formed by being cast onto blade platforms and machined to desired specifications. In such a configuration, each blade platform may be provided with a separate wing member. Hence, it should be understood that although particular structure has been illustrated and described for supporting the wing members  44 ,  144  to extend from the side of the disc  20 , 120 , as defined by either the disc structure itself or elements mounted to the side of the disc structure, other structure or additional structure for supporting the wing members  44 ,  144  may be provided. 
     While particular embodiments of the present invention have been illustrated and described, it would be obvious to those skilled in the art that various other changes and modifications can be made without departing from the spirit and scope of the invention. It is therefore intended to cover in the appended claims all such changes and modifications that are within the scope of this invention.