Abstract:
A seal arrangement between a vane assembly and a static shroud assembly reduces gas path leakage and beneficially improves gas turbine performance.

Description:
FIELD OF THE INVENTION  
       [0001]     The present invention relates to gas turbine engines, and particularly to seal means for the air leakage existing between the outer shroud of the rotor blades and adjacent stator vane shroud.  
       BACKGROUND OF THE INVENTION  
       [0002]     It is well-known to be undesirable to have uncontrolled air leakage between the shrouds of a vane ring and an adjacent turbine static shroud because leakage is a loss of energy and adverse to fuel economy.  
         [0003]     Various arrangements for sealing such leakages have been proposed, such as a continuous seal ring provided between successive shrouds. Due to the high temperature working condition of a gas turbine, the continuous seal ring requires a low thermal expansion in order to ensure an adequate seal. However, such a seal will be adversely affected when successive shrouds have different thermal expansions during engine operation. Therefore there is a need for improved seal means which will be more adequate under high temperature working conditions of gas turbine engines.  
       SUMMARY OF THE INVENTION  
       [0004]     One object of the present invention is to provide an improved seal configuration.  
         [0005]     In accordance with one aspect of the present invention, there is provided a seal assembly for minimizing fluid leakage between an end of an annular vane assembly and an end of an annular static shroud assembly of a gas turbine engine. The seal assembly comprises a primary seal comprised of co-operating abutting radial surfaces of the vane assembly and static shroud assembly and a secondary seal including a feather seal received within a cavity, the cavity being at least partially formed between two annular recesses defined in the radial abutting surfaces.  
         [0006]     In accordance with another aspect of the present invention, there is provided a turbine stator structure comprising an annular upstream shroud having a continuous circumferential downstream end, an annular downstream shroud coaxial with the upstream shroud, having a continuous circumferential upstream end abutting the downstream end of the upstream shroud to thereby provide a primary seal between the shrouds. Opposed circumferential recesses are defined in the respective abutting ends of the shrouds, thereby forming an annular cavity crossing a boundary between the abutting ends. A sealing ring is received within the cavity, abutting an annular axial surface of the cavity to substantially cover a line of the boundary on the annular axial surface.  
         [0007]     In accordance with further aspect of the present invention, there is provided a seal assembly for minimizing fluid leakage between a turbine vane assembly and a turbine static shroud assembly, the vane and shroud assemblies having planar radially-extending annular surfaces facing one another, the seal assembly comprising annular recesses defined in the respective annular surfaces, and a feather seal extending between the recesses. The feather seal preferably extends substantially around but is less than a complete circumference of the annular recesses to thereby permit interference-free circumferential thermal expansion of the feather seal.  
         [0008]     The present invention advantageously provides a simple seal configuration for minimizing a radial fluid leakage between successive shrouds without being substantially affected by thermal expansion of either the metal seal ring or the shrouds, and will provide an adequate seal even when the successive shrouds have the same or different thermal expansions. These and other advantages of the present invention will be better understood with reference to preferred embodiments of the present invention to be described hereinafter. 
     
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0009]     Reference will now be made to the accompanying drawings showing by way of illustration preferred embodiments, in which:  
         [0010]      FIG. 1  is a schematic cross-sectional view of a turbofan gas turbine engine, as an example illustrating an application of the present invention;  
         [0011]      FIG. 2  is a partial cross-sectional view of a turbine section of the engine of  FIG. 1 , showing a first embodiment of the present invention;  
         [0012]      FIG. 2A  is a cross-sectional view of the embodiment of  FIG. 2 ;  
         [0013]      FIG. 3  is a partial cross-sectional view of  FIG. 2  in an enlarged scale, showing details of the embodiment;  
         [0014]      FIG. 4  is a partial cross-sectional view similar to  FIG. 3 , showing thermal expansions during engine operation;  
         [0015]      FIG. 5  is a partial cross-sectional view similar to  FIG. 3 , showing an alternative configuration according to a second embodiment of the present invention;  
         [0016]      FIG. 6  is a partial cross-sectional view similar to  FIG. 3 , showing a further alternative configuration according to a third embodiment of the present invention;  
         [0017]      FIG. 7  is a partial cross-sectional view similar to  FIG. 3 , showing a still further alternative configuration according to a fourth embodiment of the present invention; and  
         [0018]      FIG. 8  is a partial cross-sectional view similar to  FIG. 3 , showing a still further alternative configuration according to a fifth embodiment of the present invention. 
     
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS  
       [0019]     Referring to  FIGS. 1 and 2 , a turbofan gas turbine engine incorporating an embodiment of the present invention is presented as an example of the application of the present invention, and includes a housing or a nacelle  10 , a core casing  13 , a low pressure spool assembly seen generally at  12  which includes a fan  14 , low pressure compressor  16  and low pressure turbine  18 , and a high pressure spool assembly seen generally at  20  which includes a high pressure compressor  22  and a high pressure turbine  24 . There is provided a combustor seen generally at  25  which includes an annular combustor  26  and a plurality of fuel injectors  28  for mixing liquid fuel with air and injecting the mixed fuel/air flow into the annular combustor  26  to be ignited for generating combustion gases. The low pressure turbine  18  and high pressure turbine  24  include a plurality of stator vane stages  30  and rotor stages  31 . Each of the rotor stages  31  has a plurality of rotor blades  33  encircled by a shroud assembly  32  and each of the stator vane stages  30  includes a stator vane assembly  34  which is positioned upstream and/or downstream of a rotor stage  31 , for directing combustion gases into or out of an annular gas path  36  within a corresponding shroud assembly  32 , and through the corresponding rotor stage  31 .  
         [0020]     Referring to  FIGS. 2, 2A  and  3 , a combination of the turbine shroud assembly  32  and the stator vane assembly  34  is described. The shroud assembly  32  includes a plurality of shroud segments  37  (only one shown) each of which includes a shroud ring section  38  having two radial legs  40 ,  42  with respective hooks (not indicated) conventionally supported within an annular shroud support structure (not shown) formed with a plurality of shroud support segments. The annular shroud support structure is in turn supported within the core casing  13  of  FIG. 1 . The shroud segments  37  are joined one to another in a circumferential direction and thereby form the shroud assembly  32  which encircles the rotor blades  33  and in combination with the rotor stage  31  defines a section of an annular gas path  36 . The shroud assembly  32  includes an upstream end (not indicated) and a downstream end  50 .  
         [0021]     The stator vane assembly  34  is disposed, for example, downstream of the rotor stage  31 , and includes a plurality of stator vane segments  52  (only one shown) joined one to another in a circumferential direction. The stator vane segments  52  each include an inner platform (not shown) conventionally supported on a stationary support structure (not shown) and an outer platform referred to as a stator vane shroud segment  56  to form a stator vane shroud which is conventionally supported within the annular shroud support structure. One or more (only one shown) air foils  58  radially extending between the inner platform and the stator vane shroud segment  56  divide a downstream section of the annular gas path  36  relative to the rotor stage  31 , into sectoral gas passages for directing combustion gas flow out of the rotor stage  31 .  
         [0022]     Compressed cooling air (as indicated by the arrows in  FIG. 2 ) is introduced within the shroud support structure to cool the shroud assembly  32  and the stator vane assembly  34 . The pressure of the cooling air within a cavity  60  defined between the shroud support structure and the shroud assembly  32  as well as the stator vane assembly  34 , is referred to as a “vane feed pressure” and is higher than the pressure of the combustion gas in the annular gas path  36  which is referred to as the “gas path pressure”. Therefore, it is desirable to provide a seal between the shroud assembly  32  and the stator vane shroud of the stator vane assembly  34  in order to impede cooling air flow from leaking into the gas path  36 , which causes cooling air to be wasted and thereby adversely affects engine performance efficiency and part durability.  
         [0023]     The downstream ends of the respective shroud ring section  38  in combination form the continuously circumferentially downstream end  50  of the shroud assembly  32 , preferably having a substantially flat radial surface  62  thereof. Similar to the shroud ring section  38 , the upstream ends of the respective stator vane shroud segments  56  in combination, form a continuous and circumferential upstream end  64  of the stator vane shroud of the stator vane assembly  34 , preferably having a substantially flat radial surface  66 . The substantially flat annular radial surface  62  of the shroud downstream end  50  abuts the substantially flat annular radial surface  66  of the upstream end  64  of the stator vane shroud, thereby providing a primary seal to prevent air leakage between the successive shroud assembly  32  and the stator vane assembly  34 , into the gas path  36 .  
         [0024]     Nevertheless, air leaking passages to an extent exist between the successive shroud assembly  32  and the stator vane assembly  34  through the primary seal formed by the abutting flat annular radial surfaces  62 ,  66 , due to various factors such as manufacturing tolerances, thermal expansion, etc. In order to further minimize air leakage between the successive shroud assembly  32  and the stator vane assembly  34 , a secondary seal is provided.  
         [0025]     Each of the shroud segments  37  includes a groove (not indicated) extending circumferentially from one side to the other through the downstream end thereof, thereby defining an annular recess  68  in the downstream end  50  of the shroud assembly  32  which extends from the substantially flat annular radial surface  62  into the downstream end  50 . A groove (not indicated) is also provided in each of the stator vane shroud segments  56 , extending from one side to the other through the upstream end thereof, thereby defining an annular recess  70  which extends from the substantially flat annular radial surface  66  of the upstream end  64  of the stator vane shroud of the stator vane assembly  34 . The two annular recesses  68 ,  70  are substantially aligned with each other to form an annular cavity  72 .  
         [0026]     A sealing ring  74  is received within the annular cavity  72 . The feather seal  74  in the embodiment shown in  FIGS. 2, 2A  and  4 , preferably includes a feather seal having a curved metal band having a generally rectangular cross-section loosely received within the annular cavity  72 . Therefore, under the pressure differential between the vane feed pressure in the cavity  60  and the gas path pressure in the annular gas path  36 , the seal  74  is pressed radially inwardly, (as shown by the arrows in  FIG. 3  representing the air pressure differential) to abut an annular axial surface  76  of the annular cavity  72 . Because the annular cavity  72  crosses a boundary between the abutting ends  50 ,  64  of the successive shroud assembly  32  and stator vane shroud of the stator vane assembly  34 , the seal  74  substantially covers a line of the boundary (not indicated) on the annular axial surface  76 , thereby minimizing a radial fluid leakage through those fluid leaking passages formed between the abutting ends  50 ,  64  of the successive shroud assembly  32  and stator vane shroud of the stator vane assembly  34 . Seal  74  may comprise a plurality of seal segments (not shown) circumferentially arranged, if desired.  
         [0027]     The seal  74  as shown in  FIG. 2A , includes opposed ends  78 ,  80  defining a very small gap  81  therebetween to allow for thermal expansion thereof. The small gap  81  will cause a very small air leakage therebetween, the quantity of which may be accurately determined and controlled. Nevertheless, the seal  74  preferably provides a secondary seal in addition to the primary seal formed between the abutting annular radial surfaces  62 ,  66 , and therefore the leakage through the small gap  81  is insignificant enough to be ignored. However, if desired, the seal  74  may provide a primary seal between the vane and static shroud, which will be further described below with reference to  FIG. 7 .  
         [0028]     The shroud assembly  32  has a substantially different configuration from the stator vane shroud of the stator vane assembly  34 . In the stator vane assembly  34 , the stator vane shroud segments  56  may be integrated with one or more air foils  58 . Therefore, the thermal expansion of the shroud assembly  32  may be different from that of the stator vane shroud of the stator vane segments  34  during engine operation. Furthermore, due to the different configurations, the shroud ring segments  37  and the stator vane shroud segments  56  may be fabricated in different materials which also results in different thermal expansions during engine operation. As shown in  FIG. 4 , different thermal expansions of the shroud assembly  32  and stator vane shroud of the stator vane assembly  34  will cause a radial displacement therebetween, which results in misalignment of the two annular recesses  68 ,  70 . Due to the loose accommodation of the seal  74  and the very thin cross-section thereof which results in flexibility, the seal  74  under the pressure differential as shown by the arrows, will still substantially seal the line of the boundary between the ends  50 ,  64 . In contrast to the seal  74  of the present invention, continuous seal rings used in prior art have a tendency to keep the diameter thereof equal at two sides thereof, which results in difficulties to substantially seal the line of the boundary of the abutting ends  50 ,  64  when the annular recesses  68 ,  70  are misaligned.  
         [0029]     In other embodiments described below, similar parts are identified with numerals similar to those of the description of the first embodiment and will not be redundantly described.  
         [0030]     The annular cavity and the seal of the present invention can be in various cross-sections. For example, in accordance with a second embodiment of the present invention and illustrated in  FIG. 5 , an annular cavity  72   a  is formed by two annular recesses  68   a ,  70   a  which are at angles to each other. The seal  74   a  includes a circumferentially extending seal which is angled along a central axis (not indicated) such that the two sides thereof are angled to correspond with angled orientation of the two annular recesses  68   a  and  70   a.    
         [0031]      FIG. 6  illustrates a third embodiment of the present invention in which the seal  74   b  includes a circumferentially extending seal having a curved cross-section such that the opposite sides  78 ,  80  thereof, have a diameter greater than the diameter of the middle portion therebetween.  
         [0032]      FIG. 7  illustrates a fourth embodiment of the present invention in which the seal  74   c  includes a circumferentially extending seal having two side portions  82 ,  84  curved radially outwardly with a radially outwardly arched middle portion  86 , to form a “dog bone” shaped cross-section.  
         [0033]      FIG. 8  illustrates a fifth embodiment of the present invention in which the seal  74   d , similar to the embodiment of  FIG. 7 , includes a circumferentially extending seal having opposed side portions  82 ,  84  curved preferably radially and outwardly. However, the middle portion (not indicated) between the curved side portions  82 ,  84  of the seal  74   d , is preferably generally flat, in contrast to the arched profile of the embodiment of  FIG. 7 . It is noted that the ends  50 ,  64  of the respective shroud assembly  32  and stator vane assembly do not a but one another, leaving a gap therebetween. This embodiment illustrates the applicability of the present invention when the shroud assembly  32  and stator vane assembly  34  do not provide a primary seal therebetween. In this embodiment, the seal  74   c  provides primary sealing between the adjacent turbine components.  
         [0034]     The seals  74   b ,  74   c  and  74   d  in  FIGS. 6-8  present a further aspect of the present invention. The cross-sectional dimension of the seals  74   b ,  74   c  and  74   d  is smaller in width than the annular cavity  72 , but the seals  74   b ,  74   c  and  74   d  are not loosely received within the annular cavity  72  due to the specifically profiled cross-sections thereof. When the seals  74   b ,  74   c  and  74   d  are placed within the annular cavity  72 , the opposed sides  78 ,  80  of the seal  74   b  or the opposed curved side portions  82 ,  84  of the seals  74   c  and  74   d , are compressed within the annular cavity  72 , resulting in a resilient deformation thereof which produces a radial pre-load to the seals  74   b ,  74   c  and  74   d . This radial pre-load advantageously ensures an effective seal of the seals  74   b ,  74   c  and  74   d  over the line of the boundary of the abutting ends  50 ,  64  of the successive shroud assembly  32  and the stator vane shroud of the stator vane assembly  34 , even when the pressure differential between the vane feed pressure in the cavity  60  and the gas path pressure in the annular gas path  36  of  FIG. 2  is relatively small. These pre-load types of seals  74   b ,  74   c  and  74   d  are also adapted to compensate for misalignment of the annular recesses  68 ,  70  resulting from different thermal expansions of the shroud assembly  32  and the stator vane shroud of the stator vane assembly  34 . This feature is assisted by flexible nature of the seal configuration, as disclosed above.  
         [0035]     The above-described embodiments are exemplary and are not intended to limit the present invention. Modifications and improvements to the above-described embodiments may made without departure from the principle of the present invention. For example, the seal configuration according to the present invention can be applied to any successive annular components of a gas turbine engine such as successive sections of a fan blade casing or compressor portion of a gas turbine engine. The present invention can also be applicable to gas turbine engine types other than turbofan turbine engines. Therefore the scope of the present invention is intended to be limited solely by the scope of the appended claims.