Abstract:
Disclosed are assemblies and articles for restricting leakage of a pressurized fluid from a cavity flanked by a vane support and a bladed rotor assembly. In accordance with an embodiment of the invention, the vane support defines a circumferential channel, and a interrupted rim region of the bladed rotor assembly defines a segmented ring. The segmented ring protrudes outward from the bladed rotor assembly, spans across the cavity and into the channel to define a seal.

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS  
       [0001]     This application discloses subject matter related to copending U.S. patent applications “COMBINED BLADE ATTACHMENT AND DISK LUG FLUID SEAL” (APPLICANT REFERENCE NUMBER EH-11598) and “BLADE NECK FLUID SEAL” (APPLICANT REFERENCE NUMBER EH-11507) filed concurrently herewith. 
     
    
     STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT  
       [0002]     This invention was made with Government support under F33615-98-C-2801 awarded by the United States Air Force. The Government has certain rights in this invention. 
     
    
     BACKGROUND OF THE INVENTION  
       [0003]     (1) Field of the Invention  
         [0004]     The invention relates to gas turbine engines, and more specifically to a seal for providing a fluid leakage restriction between components within such engines.  
         [0005]     (2) Description of the Related Art  
         [0006]     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 disposed in the annular duct.  
         [0007]     Since the combustion gas temperature may reach 2000 degrees Fahrenheit or more, some blade and vane stages are cooled with lower temperature cooling air for improved durability. Air for cooling the first-stage blades bypasses the combustor and is directed to an inner diameter 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 and into a series of conduits within each blade, thus providing the required cooling.  
         [0008]     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 challenging 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 inner cavities with varying degrees of success.  
         [0009]     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 is fluidly coupled to 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.  
         [0010]     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 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.  
         [0011]     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 and cooling air 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 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  
       [0012]     In accordance with an embodiment of 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 segmented ring defined by the bladed rotor assembly and a channel defined by the vane support. 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. An interrupted rim region extends radially outward from a radius circumscribing a radially innermost floor of each slot to the outermost rim. The segmented ring extends axially outward from the interrupted rim region towards the inner cavity. The circumferential channel defined by the vane support is open to the inner cavity and is located radially proximate the axially extending ring. The ring spans across the cavity and into the channel to define a seal with a more consistent leakage restriction throughout the entire range of engine operating conditions. Since a cooling air leakage restriction occurs at both inner and outer radial locations, the radial growth of the rotor assembly in relation to the vane support is accounted for.  
         [0013]     Also, by locating the seal radially outboard and in the interrupted rim region of the disk, temperature rise and parasitic drag due to pumping are minimized. Engine rotating mass is reduced with the elimination of freestanding sideplates and complex, multi-step labyrinth seal hardware as well.  
         [0014]     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 an example a seal in accordance with a preferred embodiment of the invention.  
     
    
     BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS  
       [0015]      FIG. 1  is a simplified schematic sectional view of a gas turbine engine along a central, longitudinal axis.  
         [0016]      FIG. 2  is 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.  
         [0017]      FIG. 2   a  is a detailed view of a seal in accordance with an embodiment of the present invention.  
         [0018]      FIG. 3  is a partial isometric view of the rotor assembly of  FIG. 2  showing a seal in accordance with an embodiment of the present invention.  
         [0019]      FIG. 4  is a partial front view of the rotor assembly of  FIG. 2  showing a seal in accordance with an embodiment of the present invention.  
         [0020]      FIG. 5  is a simplified sectional view of a seal in accordance with an embodiment of the present invention as assembled.  
         [0021]      FIG. 6  is a simplified sectional view of a seal in accordance with an embodiment of the present invention during an engine take-off condition.  
         [0022]      FIG. 7  is a simplified sectional view of a seal in accordance with an embodiment of the present invention during an engine cruise condition.  
     
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0023]     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 .  
         [0024]     Referring now to the example of  FIGS. 2 and 2   a,  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 comprises a disk  56  and a plurality of outwardly extending blades  58 , rotationally disposed about the central axis  11 . As best shown in  FIGS. 3 and 4 , 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 the periphery of each face  68 ,  69 .  
         [0025]     During the engine  10  operation, 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 dedicated to 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  radially flanked by the blade attachment  66  and the disk  56 . The cooling air  40  exits the blade  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 durability of the blade  58 .  
         [0026]     An exemplary seal  80  in accordance with an embodiment of the invention separates the inner cavity  50  from the annular duct  30 , thus 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, thus improving engine efficiency.  
         [0027]     The exemplary seal  80  comprises a channel  84  in the vane support  52  and a segmented ring  86  defined by the rotor assembly  54 . The channel  84  is circumferentially disposed and has a radial height  88 , an axial depth  90  and is open to the inner cavity  50 . In the example shown in  FIGS. 2 and 2   a,  the channel  84  has a ‘C’ shaped cross sectional profile; however, other cross sectional profiles may be used. The channel  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 radially inner land  94  and a radially outer land  96  are affixed to an inner radial face  98  and an outer radial face  100  of the channel  84  respectively. The lands  94 ,  96  are comprised of a honeycomb, abradable rubber or other structure known in the sealing art.  
         [0028]     The segmented ring  86  is radially located in an interrupted rim region  110  of the disk  56 . The interrupted rim region  110  extends radially outward from a radius  112  circumscribing a floor  114  of each slot  62  to the outer rim  60 . As best shown in  FIG. 3 , a first number  164  of the ring segments are defined by the disk lugs  64  and a second number  166  of the ring segments are defined by the blade attachments  66 . The first number of segments  164  are preferably formed with the disk  56  prior to milling or broaching of the slots  62 . The second number of segments  166  are preferably cast or forged integrally with the blades  58  and machined with the attachment  66 . With the blades  58  interposed with the lugs  64 , the first  164  and second  166  ring segments substantially align, defining a complete segmented ring  86 .  
         [0029]     Referring now to  FIG. 4 , tangential sealing between adjacent ring segments  164 ,  166  occurs as centrifugal forces draw the blade  58  radially outward against the lugs  64  during engine operation. To achieve this sealing, the segmented ring  86  is radially positioned to include a contact surface  168  located at the interface of the lug  64  and the attachments  66 . Although an innermost contact surface  168  is included in the example for reduced weight, any one or more of the contact surfaces  168  may be included.  
         [0030]     A circumferential runner  170  extends radially outward from the segmented ring  86  and a circumferential runner  170  extends radially inward from the segmented ring  86 . It is preferable for the axial width of the runners  170  to be as thin as possible adjacent to the lands  94 ,  96  to reduce the velocity of any cooling air  40  flowing there between. Although the runners  170  are shown in the figures at the forward extent of the segmented ring  86 , multiple runners  170  may be positioned anywhere along the axial length of the segmented ring  86 . Since intermittent contact between a runner  170  and a land  94 , or  96  may occur, a coating, hard face or other wear-resistant treatment is typically applied to the runner  170 .  
         [0031]     With the rotor assembly  54  installed in the high pressure turbine  18 , the segmented ring  86  extends outward from the interrupted rim region  110 , spans across the inner cavity  50  and into the channel  84 , aligning the runners  170  axially with the lands  94 ,  96 . The radial height  88  of the channel  84  is slightly oversized to provide sufficient clearance between the lands  94 ,  96  and the runners  170 , preventing interference while being assembled and during operation of the engine  10 . As shown in  FIG. 5 , an inner clearance C INNER  of about (0.020) inch and an outer clearance C OUTER  of about (0.020) inch ensure that the runners  170  do not interfere with the lands  94 ,  96  during assembly.  
         [0032]     By utilizing at least two radially opposed runners  170 , a more consistent leakage restriction is maintained in the seal  80  throughout all engine-operating conditions. During engine take-off conditions, as shown in  FIG. 6 , a maximum radial growth of the rotor assembly  54  occurs, closing the outer clearance C OUTER  to about (0.000) inch and opening the inner clearance C INNER  to about (0.040) inch. During engine cruise conditions, as shown in  FIG. 7 , the radial growth of the rotor assembly  54  stabilizes and the outer clearance C OUTER  is about (0.005) inch while the inner clearance C INNER  is about (0.035) inch.  
         [0033]     Although an exemplary seal  80  has been 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.  
         [0034]     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.