CROSS-REFERENCE TO RELATED APPLICATIONS  
       [0001]     This application discloses subject matter related to copending US patent applications “HAMMERHEAD FLUID SEAL” (APPLICANT REFERENCE NUMBER EH-11279) and “COMBINED BLADE ATTACHMENT AND DISK LUG FLUID SEAL” (APPLICANT REFERENCE NUMBER EH-11598) each filed concurrently herewith. 
     
    
     BACKGROUND OF THE INVENTION  
       [0002]     (1) Field of the Invention  
         [0003]     The invention relates to gas turbine engines, and more specifically to a seal for providing a fluid leakage restriction between components within such engines.  
         [0004]     (2) Description of the Related Art  
         [0005]     Gas turbine engines operate by burning a combustible fuel-air mixture in a combustor and converting the energy of combustion into a propulsive force. Combustion gases are directed axially rearward from the combustor through an annular duct, interacting with a plurality of turbine blade stages disposed within the duct. The blades transfer the combustion gas energy to one or more blades mounted on disks, rotationally disposed about a central, longitudinal axis of the engine. In a typical turbine section, there are multiple, alternating stages of stationary vanes and rotating blades in the annular duct.  
         [0006]     Since the combustion gas temperature may reach 2000 degrees Fahrenheit or more, some blade and vane stages are cooled with a lower temperature cooling air for improved durability. Air for cooling the first-stage blades bypasses the combustor and is directed to an inner cavity located between a first-stage vane support and a first-stage rotor assembly. The rotational force of the rotor assembly pumps the cooling air radially outward into a series of conduits within each blade, thus providing the required cooling.  
         [0007]     Since the outboard radius of the inner cavity is adjacent to the annular duct carrying the combustion gasses, it must be sealed to prevent leakage of the pressurized cooling air into the combustion gas stream. This area of the inner cavity is particularly difficult to seal due to the differences in thermal and centrifugal growth between the stationary, first-stage vane support and the rotating, first-stage rotor assembly. In the past, designers have attempted to seal the outboard radius of the inner cavity with varying degrees of success.  
         [0008]     An example of such an outboard radius seal is a labyrinth seal. In a typical configuration, a multi-step labyrinth seal separates the inner cavity into two regions of approximately equal size, an inner region and an outer region. Cooling air in the inner region is pumped between the rotating disk and labyrinth seal into the hollow conduits of the blades while the outer region communicates with the annular duct carrying the combustion gases. A labyrinth seal&#39;s lands must be pre-grooved to prevent interference between the knife-edge teeth and the lands during a maximum radial excursion of the rotor. By designing the labyrinth seal for the maximum radial excursion of the rotor assembly, the leakage restriction capability is reduced during low to intermediate radial excursions of the rotor assembly. Any cooling air that leaks by the labyrinth seal is pumped through the outer region and into the annular duct by the rotating disk. This pumping action increases the temperature of the disk in the area of the blades and creates parasitic drag, which reduces overall turbine efficiency. The rotating knife-edges also add additional rotational mass to the gas turbine engine, which further reduces engine efficiency.  
         [0009]     Another example of such an outboard radius seal is a brush seal. In a typical configuration, a brush seal separates the inner cavity into two regions, an inner region and a smaller, outer region. A freestanding sideplate assembly defines a disk cavity, which is in fluid communication with the inner region. Cooling air in the inner region enters the disk cavity and is pumped between the rotating sideplate and disk to the hollow conduits of the blades. The seal&#39;s bristle to land contact pressure increases during the maximum radial excursions of the rotor and may cause the bristles to deflect and ‘set’ over time, reducing the leakage restriction capability during low to intermediate rotor excursions. Any cooling air that leaks by the brush seal is pumped into the outer region by the rotating disk. This centrifugal pumping action increases the temperature of the disk in the area of the blades and creates parasitic drag, which reduces overall turbine efficiency. The freestanding sideplate and minidisk also adds rotational mass to the gas turbine engine, which further reduces engine efficiency.  
         [0010]     Although each of the above mentioned seal configurations restrict leakage of cooling air under certain engine operating conditions, a consistent leakage restriction is not maintained throughout all the radial excursions of the rotor. The seals may also increase the temperature of the disk due to centrifugal pumping, reduce engine efficiency due to parasitic drag and add additional engine weight. What is needed is a seal that maintains a more consistent fluid leakage restriction throughout all the radial excursions of the rotor, without negatively affecting disk and cooling air temperature, engine efficiency or engine weight.  
       BRIEF SUMMARY OF THE INVENTION  
       [0011]     In accordance with the present invention, there is provided a seal for restricting leakage of pressurized cooling air from an inner cavity flanked by a vane support and a bladed rotor assembly. The seal comprises a land defined by the vane support and a segmented ring defined by the bladed rotor assembly. The bladed rotor assembly includes a disk rotationally disposed about a central axis of the engine. The disk includes a radially outermost rim and a plurality of slots circumferentially spaced about the rim for accepting an equal plurality of blades. Each blade contains a radially lowermost attachment, which engages a slot in a sliding arrangement. A neck region extends outboard of the rim from the attachment to a platform of each blade. A segmented ring extends from the neck region to define a segregated inner and outer cavity. The land defined by the vane support is located radially above the inner cavity, proximate to the segmented ring. The segmented ring spans across the inner cavity, interacting with the land to define the seal.  
         [0012]     By locating the seal radially outboard and in the neck region of the blades, temperature rise and parasitic drag due to tangential on board injector (TOBI) placement and pumping are minimized. Also, engine rotating mass is reduced with the elimination of freestanding sideplates and complex, multi-step labyrinth seal hardware as well.  
         [0013]     Other features and advantages will be apparent from the following more detailed descriptions, taken in conjunction with the accompanying drawings, which illustrate by way of examples a seal in accordance with specific embodiments of the invention.  
     
    
     BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS  
       [0014]      FIG. 1  illustrates a simplified schematic sectional view of a gas turbine engine along a central, longitudinal axis.  
         [0015]      FIG. 2  illustrates a partial sectional view of a turbine rotor assembly of the type used in the engine of  FIG. 1 , showing a seal in accordance with an embodiment of the present invention.  
         [0016]      FIG. 3  illustrates a partial sectional view of a turbine rotor assembly of the type used in the engine of  FIG. 1 , showing a multiple step seal in accordance with an embodiment of the present invention.  
         [0017]      FIG. 4  illustrates a partial isometric view of the turbine rotor assembly of  FIG. 2 .  
         [0018]      FIG. 5  illustrates a partial front view of the turbine rotor assembly of  FIG. 2 .  
         [0019]      FIGS. 6   a - 6   h  illustrates a series of enlarged schematics illustrating various seals of  FIGS. 2 and 3  in accordance with several embodiments of the present invention.  
     
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0020]     The major sections of a typical gas turbine engine  10  of  FIG. 1  include in series, from front to rear and disposed about a central longitudinal axis  11 , a low-pressure compressor  12 , a high-pressure compressor  14 , a combustor  16 , a high-pressure turbine  18  and a low-pressure turbine  20 . A working fluid  22  is directed rearward through the compressors  12 ,  14  and into the combustor  16 , where fuel is injected and the mixture is burned. Hot combustion gases  24  exit the combustor  16  and expand within an annular duct  30  through the turbines  18 ,  20  and exit the engine  10  as a propulsive thrust. A portion of the working fluid  22  exiting the high-pressure compressor  14 , bypasses the combustor  16  and is directed to the high-pressure turbine  18  for use as cooling air  40 .  
         [0021]     Referring now to  FIGS. 2 and 3 , an inner cavity  50  is located radially inward of the annular duct  30  and axially between a first-stage vane support  52  and a first-stage rotor assembly  54 . The rotor assembly  54  comprises a disk  56  and a plurality of outwardly extending blades  58  rotationally disposed about the central axis  11 . As best shown in  FIGS. 4 and 5 , the disk  56  includes a radially outermost rim  60 , a plurality of fir tree profiled slots  62 , and a plurality of lugs  64  alternating with the slots  62  about the circumference of the rim  60 . Each slot  62  accepts a radially lower most attachment  66  of a blade  58  in a sliding arrangement. One or more teeth  67  extend between a forward, axial face  68  and a rearward, axial face  69  of the attachment  66 , engaging adjacent lugs  64  to prevent loss of the blade  58  as the disk  56  rotates. The one or more teeth  67 , project a complementary fir tree profile about a periphery of each face  68 ,  69 .  
         [0022]     During the operation of the engine  10 , pressurized cooling air  40  is pumped into the inner cavity  50  by a duct  70 , where a major portion of the cooling air  40  is used for internally cooling the blades  58 . The cooling air  40  enters the blades  58  via a series of radially extending conduits  72  communicating with a plenum  74  flanked by the blade attachment  66  and the disk  56 . The cooling air  40  exits the blades  58  via a series of film holes  76 . To ensure a continuous flow of cooling air  40  through the blade  58 , the pressure of the cooling air  40  must remain greater than the pressure of the combustion gases  24  or the combustion gases  24  may backflow into the film holes  76 , potentially affecting the blade  58  durability.  
         [0023]     An exemplary seal  80  in accordance with an embodiment of the invention separates the inner cavity  50  from the annular duct  30 , ensuring adequate cooling air  40  pressure throughout all engine-operating conditions. The seal  80  is located radially inward of the annular duct  30 , defining an outer cavity  82  therebetween. Since the outer cavity  82  is relatively small, any leakage of cooling air  40  through the seal  80  is subject to relatively minimal pumping by the rotor assembly  54  prior to mixing with the combustion gases  24 . This level of pumping has limited negative impact on disk  56  temperature and aerodynamic drag, which in turn, improves engine-operating efficiency.  
         [0024]     The exemplary seals  80  of  FIGS. 2 and 3 , comprise a circumferentially disposed land  84  defined by the vane support  52  and a segmented ring  86  defined by the rotor assembly  54 . In the examples shown, the lands  84  have a linear cross sectional profile; however, other profiles such as those shown in the examples of  FIGS. 6a-6h  may also be used. Lands  84  at differing radial locations provide an increased restriction over a single land  84 . A land  84  may be integrally defined by the vane support  52  or may be defined by a separate arm  92  and affixed to the vane support  52  by welding, bolting, riveting or other suitable means. A land  84  is generally affixed to faces  94  of the vane support  52  or arm  92  by brazing and is comprised of honeycomb or any other abradable structure known in the sealing art.  
         [0025]     The segmented ring  86  is radially located in a neck region  96  of the blades  58 . The neck region  96  extends radially outward, above the rim  60 , from the attachment  66  to a platform  98  that supports an airfoil  100  and defines the inner radial contour of the annular duct  30 . Individual ring segments  186  extend axially outward from the neck region  96  of each blade  58  and are formed by casting, turning, grinding, broaching, electrodischarge (EDM) or other suitable process. With the blades  58  interposed with the lugs  64 , adjacent ring segments  186  substantially align, defining a complete segmented ring  86 . A single segmented ring  86 , as shown in  FIG. 2 , may be used, or multiple segmented rings  86 , as shown in  FIG. 3 , may also be used. The addition of multiple segmented rings  86  provides a greater leakage restriction, but the actual number may be limited by space and weight requirements.  
         [0026]     A runner  200 , also know as a knife edge, extends outward from a segmented ring  86  as shown in  FIGS. 2 and 3 . The addition of multiple runners  200  provides a greater cooling air  40  leakage restriction, but the actual number may be limited by space and weight requirements. The width of a runner  200  should be as thin as possible, adjacent to a land  84 , to reduce the velocity of any cooling air  40  flowing there between. Since intermittent contact between a runner  200  and a land  84  may occur, a coating, hardface or other wear-resistant treatment is typically applied to the runners  200 . A runner  200  may also be canted from between about 22.5 degrees to about 68 degrees, preferably 55 degrees, relative to the engine axis  11 . By canting a runner  200  in the direction opposing the cooling air  40  flow, a damming effect is created, providing for an increased leakage restriction. Canting a runner  200  also reduces the length of the thicker, segmented ring  86 , reducing weight even further. Several examples of a runner  200  are shown in  FIGS. 6   a - 6   h.    
         [0027]     Referring now to  FIGS. 4 and 5 , cooling air  40  leakage between adjacent ring segments  186  may be minimized by utilizing localized sealing means. In an exemplary embodiment, sealing between adjacent ring segments  186  is achieved with a matched tongue  190  and groove  192  joint, located at the interface of adjacent ring segments  186 . Although the example shows a linear tongue  190  and groove  192  joint, any suitable shaped joint may be used. It is to be understood that other sealing means known in the art such as feather seals, shiplap seals and the like may also be used.  
         [0028]     With the rotor assembly  54  installed in the high pressure turbine  18  as shown in  FIGS. 2 and 3 , a segmented ring  86  extends outward from the neck region  96  of the blades  58 , spans across the inner cavity  50 , aligning a runner  200  axially with a land  84 . Sufficient clearance between a runner  200  and a land  84  prevents interference during assembly and during engine  10  operation.  
         [0029]     Although an exemplary seal  80  is shown positioned between a stationary member and a rotating member, it is to be understood that an exemplary seal  80  may also be located between two rotating members or two stationary members as well.  
         [0030]     While the present invention has been described in the context of specific embodiments thereof, other alternatives, modifications and variations will become apparent to those skilled in the art having read the foregoing description. Accordingly, it is intended to embrace those alternatives, modifications and variations as fall within the broad scope of the appended claims.
Disclosed are assemblies and articles for restricting leakage of a pressurized fluid from a cavity. In accordance with an embodiment of the invention, a vane support defines a land, and a neck region of a bladed rotor assembly defines a segmented ring. The segmented ring protrudes outward from the bladed rotor assembly in the neck region, spans across the cavity and cooperates with the land to define a seal.