Patent Abstract:
Disclosed are various embodiments of a gas turbine blade to vane interface seal for restricting leakage of cooling air and improving the apportioning of the cooling air to the seal. Accordingly, a turbine rotor contains a first and a second stage of radially extending and circumferentially distributed blades. The stages are separated axially from one another by an annular coupling located radially inboard of the blades, forming a chamber therebetween. Interposed between the blade stages is a vane stage. The vane stage contains a land, facing radially inwardly. A ring projects axially from each of the first and second blade stages towards the vane stage. A ring may also project radially from the coupling towards the vane stage. The rings radially cooperate with the land and together form the blade to vane interface seal. The coupling contains an aperture for radially introducing a cooled fluid to the chamber for use in cooling the seal.

Full Description:
CROSS-REFERENCE TO RELATED APPLICATIONS  
       [0001]     This application discloses subject matter related to copending U.S. patent applications “HAMMERHEAD FLUID SEAL” (Ser. No. 11/146,801), “COMBINED BLADE ATTACHMENT AND DISK LUG FLUID SEAL” (Ser. No. 11/146,798) and “BLADE NECK FLUID SEAL” (Ser. No. 11/146,660), each filed on Jul. 7, 2005. 
     
    
     BACKGROUND OF THE INVENTION  
       [0002]     (1) Field of the Invention  
         [0003]     The invention relates to gas turbine engines, and more specifically to a cooled fluid sealing arrangement disposed between blades and vanes of such engines.  
         [0004]     (2) Description of the Related Art  
         [0005]     Gas turbine engines operate by compressing ambient air with a forward compressor, injecting a fuel, burning the air-fuel mixture in a central combustor and converting the energy of combustion into a propulsive force. Combustion gases exit the combustor through an annular duct, where the gases drive one or more axial stages of circumferentially distributed turbine blades. Each bladed stage transfers the combustion gas energy to a rotor attached to a central, longitudinal shaft. Interposed with the rotating blade stages are stationary vane stages affixed to radially outer casing structures, circumscribing the rotor. Two or more rotors may operate independently of one another and at differing speeds via concentric shafts. Gas turbine engines are flexible power plants that are typically used for powering aircraft, ships and generators.  
         [0006]     In order to withstand combustion gas temperatures that regularly exceed 2000 degrees Fahrenheit and pressures exceeding 400 pounds per square inch absolute, turbine components such as blades, vanes and seals are cooled with lower-temperature, higher-pressure cooling air. The cooling air is bled from the compressors, then directed axially rearward and radially inward of the rotors to the turbine components, bypassing the combustor altogether. Once delivered to the turbine, a significant portion of the cooling air is directed radially outward to the blades, vanes and seals by the centrifugal force of the turning rotors. In order to achieve the greatest heat reduction benefit from the cooling air, the interfaces of the rotating blade stages and stationary vane stages must be effectively sealed.  
         [0007]     The interfaces of the rotating blade stages and stationary vane stages are particularly difficult to seal due to the differences in thermal and centrifugal growth between the rotors and the cases. The high relative speeds, extremely high temperatures and pressures also present seal design challenges in the turbines. In the past, designers have attempted to seal the interfaces of the rotating blade stages and stationary vane stages with varying degrees of success.  
         [0008]     An example of such a turbine seal is a labyrinth seal. In a typical blade to vane interface, a multi-step labyrinth seal, comprising stationary lands and rotating runners or knife-edges, restricts leakage of the cooling air radially outward, into the combustion gases. The runners project from annular supports, which are typically fastened to the rotor with bolted flanges and/or with snap fits. The supports are independent components, adding to the manufacturing costs and complexity of the turbine. The supports also contribute additional rotational mass to the rotors, which reduces the engine-operating efficiency. Also, the attachments at the interfaces of the supports and the rotors create an additional leakage path for the cooling air. Placement of the supports is influenced by adjacent components and typically does not optimize the distribution of the cooling air.  
         [0009]     What is needed is a blade to vane interface seal that doesn&#39;t require separate seal support components, and also improves the apportioning of cooling air to the seal itself.  
       BRIEF SUMMARY OF THE INVENTION  
       [0010]     In accordance with the present invention, there are provided rotor to stator interface seals for restricting leakage of cooling air and improving the apportioning of the cooling air to the seals.  
         [0011]     Accordingly, a turbine rotor contains a first and a second stage of circumferentially distributed blades. The blade stages are separated axially from one another by an annular coupling located radially inboard of the blades, forming a chamber therebetween. Interposed between the blade stages is a stationary vane stage. The vane stage contains a land, facing radially inwardly. A ring projects axially from each of the first and second blade stages towards the vane stage. The rings radially cooperate with the land and together form the blade to vane interface seal. The coupling contains an aperture for radially introducing cooling air to the chamber for use in cooling the seal.  
         [0012]     In another embodiment of an interface seal in accordance with the present invention, a turbine rotor contains a first and a second stage of circumferentially distributed blades. The blade stages are separated axially from one another by an annular coupling located radially inboard of the blades, forming a chamber therebetween. Interposed between the blade stages is a stationary vane stage. The vane stage contains a radially inwardly facing land. A ring projects axially from blade stages towards the vane stage. The rings radially cooperate with the land. The coupling contains an integral ring projecting radially outward and radially cooperating with the land. Together, the cooperating rings and land form the blade to vane interface seal. The coupling also contains an aperture for radially introducing cooling air to the chamber for use in cooling the seal. Although the aperture may be located anywhere along the axial length of the coupling, it is typically located forward of the vane stage.  
         [0013]     Since the sealing rings are integral with the existing blades and couplings of the gas turbine engine, separate supports are not needed and are therefore eliminated. The elimination of separate supports reduces the rotational mass of the rotors, thus improving engine-operating efficiency. Also, by relocating the rings to the blades, cooling air leakage paths are eliminated and the cooling air apportioning to the seal is improved.  
         [0014]     Other details of the present invention, as well as other objects and advantages attendant thereto, are set forth in the following detailed description and the accompanying drawings wherein like reference numerals depict like elements.  
     
    
     BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS  
       [0015]      FIG. 1  illustrates a simplified schematic sectional view of a gas turbine engine along a central, longitudinal axis.  
         [0016]      FIG. 2  illustrates a partial sectional view of a low-pressure turbine of the type used in the engine of  FIG. 1 .  
         [0017]      FIG. 3  illustrates a detailed sectional view of a blade to vane interface seal embodiment of the type used in the turbine of  FIG. 2 .  
         [0018]      FIG. 4  illustrates a detailed sectional view of another blade to vane interface seal embodiment of the type used in the turbine of  FIG. 2 .  
         [0019]      FIG. 5  illustrates an isometric view of a turbine blade of the type used in the turbine of  FIG. 2 .  
         [0020]      FIG. 6   a  illustrates a front view of a ring segment interface comprising a single chamfered edge.  
         [0021]      FIG. 6   b  illustrates a front view of a ring segment interface comprising double chamfered edges.  
         [0022]      FIG. 6   c  illustrates a front view of a ring segment interface comprising a single sloped edge.  
         [0023]      FIG. 6   d  illustrates a front view of a ring segment interface comprising dual sloped edges.  
         [0024]      FIG. 6   e  illustrates a front view of a ring segment interface comprising tangentially sloped wings.  
         [0025]      FIG. 6   f  illustrates a front view of a ring segment interface comprising a single downstream dam.  
         [0026]      FIG. 6   g  illustrates a front view of a ring segment interface comprising dual dams. 
     
    
     DETAILED DESCRIPTION OF THE INVENTION  
       [0027]     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  26 , driving the turbines  18 ,  20 . The turbines  18 ,  20 , in turn drive coupled compressors  14 ,  12  via concentric shafts  28 ,  30 , forming a high rotor spool  32  and a low rotor spool  34  respectively. Although a dual spool engine  10  is depicted in the figure, three spool engines  10  are not uncommon. The combustion gases exit the engine  10  as a propulsive thrust  36 , used to power an aircraft or a free turbine. A portion of the working fluid  22  is bled from the compressors  12 ,  14  and is directed radially inward of the combustor  16  and axially rearward to the turbines  18 ,  20  for use as cooling air  38 .  
         [0028]     In an exemplary low-pressure turbine  20  of  FIGS. 2-4 , the combustion gases  24  are directed rearward through an annular duct  40  approximately defined by a radially outer flow path  42  and a radially inner flow path  44 . Disposed circumferentially within the annular duct  40  are alternating stages of rotating blades  50   a - 50   e  and stationary vanes  52   a - 52   d . The blades  50  extend radially outward from a rotor disk  54  by roots  56  disposed radially inward of platforms  58 . Each blade  50  further comprises an airfoil  60 , extending radially between the platform  58  and an outer tip shroud  62 . The airfoil  60  has a forward facing leading edge and a rearward facing trailing edge. In some instances, the blades  50  are removable from the disks  54  and in some instances non-removable. The vanes  52  are cantilevered inward from a case  64  by hooks  66  extending radially outward from the outer tip shrouds  62 . Each vane  52  comprises an airfoil  60  that extends radially between an inner shroud  68  and an outer shroud  70 .  
         [0029]     Outer seals  72  restrict leakage of the combustion gases  74  at the outer flow path  42 . The outer seals  72  are disposed at the interface of the rotating blades  50  and the stationary case  64 . The tip shrouds  62  contain outwardly extending runners  74  that radially cooperate with inwardly facing lands  76  affixed to the case  64  by supports  78 . The radial cooperation of the runners  74  and the lands  76 , along with the rotation of the blades  50 , cause a damming effect and thus restricts leakage of the combustion gases  24  from the outer flow path  42 . Overlapping platforms  58  and a constant supply of higher pressure cooling air  38  restrict leakage of the combustion gases  24  at the inner flow path  44 .  
         [0030]     Cooling air  38 , bled from the compressors  12 , 14  is directed to bore cavities  80 . The bore cavities  80  are bounded axially by adjacent disk bores  82  and radially outwardly by an annular coupling  84 . The coupling  84  joins adjacent disks  54  with bolts, rivets, welds, threads, splines, tapers, snap fits, or other means. The coupling  84  may also be integrally formed with each of the adjacent disks  82  (not shown). The cooling air  38  is pumped radially outward, against the couplings  84 , by the rotation of the disks  54 . Apertures  86  in the couplings  84  direct the cooling air  38  into rim cavities  88 . The apertures may be circular holes, slots, or other forms and are typically, evenly distributed cirumferentially about the coupling  84 . The apertures  86  are sized to allow the appropriate cooling air  38  volume to enter the rim cavity  88 .  
         [0031]     The cooling air  38  inside the rim cavity  88  is maintained at a higher pressure than the combustion gases  24  in the annular duct  40  under all engine-operating conditions. The higher pressure cooling air  38  prevents combustion gas  24  ingestion into the rim cavities  88  and provides cooling for the blade  50  to vane  52  interface. A portion of the cooling air  38  is directed axially rearward through a plurality of slots  90  disposed between the blade roots  56  and the disk  54 . This portion of cooling air  38  reduces the temperature of the blade root  56  to disk  54  interface before being directed axially rearward to a downstream rim cavity  88 . Another portion of the cooling air  38  is directed radially outward to cool the blade  50  to vane  52  interface region.  
         [0032]     As specifically illustrated in  FIGS. 3 and 4 , seals  92  according to various embodiments of the current invention restrict the leakage of the cooling air  38  at the interfaces of the blades  50  and vanes  52 . The blade platforms  58  form one or more circumferentially segmented rings  94  that radially cooperate with inwardly facing lands  96  affixed to the vanes  52 . Also, one or more integral rings  94  may project radially outward from coupling  84  anywhere along its axial length as specifically illustrated in  FIG. 4 . The cooperation of the integral rings  94  and lands  96  form intermediate seals, which partition cavity  88  into two or more smaller cavities  88 . The radially outward projecting ring  94  is not segmented and also radially cooperates with a land  96  affixed to a vane  52 . The proximate radial position of the rings  94  and the lands  96 , along with the rotation of the blades  50 , cause a damming effect and thus restrict leakage of the cooling air  38  from the rim cavity  88 .  
         [0033]     The lands  96  may have a constant radial profile or may be stepped radially to further prevent ingestion of the combustion gases  24  into the rim cavity  88 . A land  96  may be affixed directly to the vane  52  by brazing, welding or other suitable means or may be affixed to a support  97  projecting radially inwardly from the vane  52 . The support  97  may be integrated with the vane  52  or may be affixed by brazing, welding or other suitable means. A land  96  is typically comprised of a honeycomb shaped, sheet metal structure, or any other structure and material known in the sealing art to restrict leakage.  
         [0034]     The rings  94  project axially from a platform  58  of a blade  50  in a leading edge direction, a trailing edge direction, or both directions. An integral ring  94  may also project radially from coupling  84 . With the blades  50  assembled into a disk  54 , individual ring  94  segments axially and radially align, to form a substantially complete ring  94  about central axis  11 . A ring  94  may contain one or more radially extending runners  98 , which are also known as knife-edges. The addition of multiple runners  98  provides a greater cooling air  38  leakage restriction, but the actual number may be dictated by space and/or weight limitations. The width of a runner  98  should be as thin as possible, adjacent to a land  96 , to reduce the velocity of any cooling air  38  flowing therebetween. Since intermittent contact between a runner  98  and a land  96  may occur, a coating, hardface or other wear-resistant treatment is typically applied to the runners  98 . A runner  98  may also be canted at an angle (●) from between about 22.5 degrees to about 68 degrees, preferably 55 degrees, relative to the longitudinal axis of the segmented ring  94 . By canting the runner  98  in the direction opposing the cooling air  38  flow, a damming effect is created, providing for an increased leakage restriction. Canting a runner  98  also reduces the length of the thicker, segmented ring  94 , reducing weight even further. The rings  94  and runners  98  are formed by casting, conventional machining, electrodischarge machining, chemical milling, or any other suitable manufacturing methods.  
         [0035]     As further illustrated by the blade  50  embodiment of  FIG. 5 , adjacent ring  94  segments may contain mechanical sealing elements to reduce leakage of cooling air  38  therebetween. With the blades  50  installed, a tongue  100  and a groove  102  cooperate between adjacent ring  94  segments to reduce leakage of the cooling air  38 . It is noted that the tongue  100  may be inclined radially outward to ensure it completely contacts the groove  102  under centrifugal loading. Since an increased radial thickness of the ring  94  segment is only required to accommodate the tongue  100  and groove  102 , one or more pockets  104  are typically located between the tongue  100  and groove  102  to reduce the rotational mass of the blade  50 . The pockets  104  are formed by casting, conventional machining, electrodischarge machining, chemical milling or any other suitable manufacturing methods.  
         [0036]     As illustrated in the ring  94  segment embodiments of  FIGS. 6   a - 6   g , adjacent ring  94  segments may contain aerodynamic sealing means to reduce leakage of cooling air  38  therebetween. By directing a volume of cooling air  38  and combustion gases  24  radially inward through the mechanism of reverse inward pumping, the radially outward leakage of cooling air  38  from the rim cavity  88  is opposed, and therefore reduced. In each of the figures, the reference rotation of the blades  50  is in the clockwise direction. If the rotation of the blades  50  is in the counterclockwise direction, the inventive aerodynamic sealing elements are mirrored about a plane extending through the longitudinal axis  11  of the engine  10 . Also, the upstream ring  194  segment is illustrated to the right and the downstream ring  294  segment is illustrated to the left in each of the figures.  
         [0037]      FIG. 6   a  illustrates a chamfered edge  106 , reverse pumping element. The chamfered edge  106  is located at the intersection of a tangentially facing surface  108  and a radially outer surface  110  of the upstream ring  194  segment. A volume of cooling air  38  and combustion gases  24  encounters the chamfered edge  106  and is pumped radially inward, between adjacent ring  194 ,  294  segments, by the rotation of the blades  50 . The inward pumping opposes the radially outward leakage of cooling air  38 .  
         [0038]      FIG. 6   b  illustrates a double chamfered edge  106 , reverse pumping element. A chamfered edge  106  is located at the intersection of a tangentially facing surface  108  and a radially outer surface  110  of the upstream ring  194  segment. Also, a chamfered edge  106  is located at the intersection of a tangentially facing surface  108  and a radially inner surface  112  of the downstream ring  294  segment. A volume of cooling air  38  and combustion gases  24  encounters the chamfered edges  106  and is pumped radially inward, between adjacent ring  194 ,  294  segments, by the rotation of the blades  50 . The inward pumping opposes the radially outward leakage of cooling air  38 .  
         [0039]      FIG. 6   c  illustrates a single sloped edge  114 , reverse pumping element. A sloped edge  114  is located between a radially outer surface  110  and a radially inner surface  112  of the upstream ring  194  segment. A volume of cooling air  38  and combustion gases  24  encounters the sloped edge  114  and is pumped radially inward, between adjacent ring  194 ,  294  segments, by the rotation of the blades  50 . The inward pumping opposes the radially outward leakage of cooling air  38 .  
         [0040]      FIG. 6   d  illustrates a dual sloped edge  114 , reverse pumping element. A sloped edge  114  is located between a radially outer surface  110  and a radially inner surface  112  of the upstream ring  194  segment. Also, a sloped edge  114  is located between a radially outer surface  110  and a radially inner surface  112  of the downstream ring  194  segment. A volume of cooling air  38  and combustion gases  24  encounters the sloped edges  114  and is pumped radially inward, between adjacent ring  194 ,  294  segments, by the rotation of the blades  50 . The inward pumping opposes the radially outward leakage of cooling air  38 .  
         [0041]      FIG. 6   e  illustrates a dual tangentially sloped wing  116 , reverse pumping element. A radially inner sloped wing  116  is located adjacent the tangentially facing surface  108  of the upstream ring  194  segment. Also, a radially outer sloped wing  116  is located adjacent the tangentially facing surface  108  of the downstream ring  294  segment. A volume of cooling air  38  and combustion gases  24  encounters the wings  116  and is pumped radially inward, between adjacent ring  194 ,  294  segments, by the rotation of the blades  50 . The inward pumping opposes the radially outward leakage of cooling air  38 .  
         [0042]      FIG. 6   f  illustrates a single downstream dam  118 , reverse pumping element. The tangentially facing surface  108  of the downstream ring  294  segment is radially thickened and protrudes radially outward, beyond the tangentially facing surface  108  of the upstream ring  194  segment to form the dam  118 . A volume of cooling air  38  and combustion gases  24  encounters the dam  118  and is pumped radially inward, between adjacent ring  194 ,  294  segments, by the rotation of the blades  50 . The inward pumping opposes the radially outward leakage of cooling air  38 .  
         [0043]      FIG. 6   g  illustrates a dual dam  118 , reverse pumping feature. The tangentially facing surface  108  of the downstream ring  294  segment is radially thickened and protrudes radially outward, beyond the tangentially facing surface  108  of the upstream ring  194  segment. Also, the tangentially facing surface  108  of the upstream ring  194  segment is radially thickened and protrudes radially inward, beyond the tangentially facing surface  108  of the downstream ring  294  segment. A volume of cooling air  38  and combustion gases  24  encounters the dam and is pumped radially inward, between adjacent ring  194 ,  294  segments, by the rotation of the blades  50 . The inward pumping opposes the radially outward leakage of cooling air  38 .  
         [0044]     Although a low-pressure turbine  20  is illustrated throughout the figures for succinctness, it is understood that high-pressure and mid-pressure turbines are similarly constructed and would therefore benefit from the exemplary seals  92  and rim cavity  88  cooling arrangements.  
         [0045]     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.

Technology Classification (CPC): 5